Scope Title

Photovoltaic Energy Conversion

Scope Description

Photovoltaic cell and blanket technologies that lead to significant improvements in overall solar array performance by increasing photovoltaic cell efficiency greater than 30%, increasing array mass specific power greater than 300W/ kg, decreased stowed volume, reduced initial and recurring costs, long- term operation in radiation environments, high power arrays and a wide range of space environmental operating conditions are solicited.

Photovoltaic Energy Conversion: advances in, but not limited to, the following: (1) Photovoltaic cell and blanket technologies capable of low intensity, low-temperature operation applicable to outer planetary (low solar intensity) missions, (2) Photovoltaic cell, and blanket technologies that enhance and extend performance in lunar applications including orbital, surface and transfer, (3) Solar arrays to support Extreme Environments Solar Power type missions, including long-lived, radiation tolerant, cell and blanket technologies applicable to Jupiter missions, and (4) Lightweight solar array technologies applicable to science missions using solar electric propulsion.

Current missions being studied require solar arrays that provide 1 to 20 kilowatts of power at 1 AU, greater than 300 watts/kilogram specific power, operation in the range of 0.7 to 3 AU, low stowed volume, and the ability to provide operational array voltages up to 300 volts to enable direct drive electric propulsion systems for science missions.

References

Solar Power Technologies for Future Planetary Science Missions, found at: https://solarsystem.nasa.gov/resources/548/solar-power-technologies-for-future-planetary-science-missions/

 

Scope Title

Photovoltaic Energy Conversion

Scope Description

Photovoltaic cell and blanket technologies that lead to significant improvements in overall solar array performance by increasing photovoltaic cell efficiency greater than 30%, increasing array mass specific power greater than 300W/ kg, decreased stowed volume, reduced initial and recurring costs, long- term operation in radiation environments, high power arrays and a wide range of space environmental operating conditions are solicited.

Photovoltaic Energy Conversion: advances in, but not limited to, the following: (1) Photovoltaic cell and blanket technologies capable of low intensity, low-temperature operation applicable to outer planetary (low solar intensity) missions, (2)Photovoltaic cell, and blanket technologies that enhance and extend performance in lunar applications including orbital, surface and transfer, (3) Solar arrays to support Extreme Environments Solar Power type missions, including long-lived, radiation tolerant, cell and blanket technologies applicable to Jupiter missions, and (4) Lightweight solar array technologies applicable to science missions using solar electric propulsion.

Current missions being studied require solar arrays that provide 1 to 20 kilowatts of power at 1 AU, greater than 300 watts/kilogram specific power, operation in the range of 0.7 to 3 AU, low stowed volume, and the ability to provide operational array voltages up to 300 volts to enable direct drive electric propulsion systems for science missions.

References

Solar Power Technologies for Future Planetary Science Missions, found at:
https://solarsystem.nasa.gov/resources/548/solar-power-technologies-for-future-planetary-science-missions/

NASA outlines New Lunar Science, Human Exploration Missions, found at:
https://www.nasa.gov/feature/nasa-outlines-new-lunar-science-human-exploration-missions

NASA Science Missions, found at:
https://science.nasa.gov/missions-page?field_division_tid=All&field_phase_tid=3951

Expected TRL or TRL range at completion of the project 3 to 5

Desired Deliverables of Phase II

Prototype, Analysis, Hardware, Research

Desired Deliverables Description

Phase I deliverables include detailed reports with proof- of- concept and key metrics of components tested and verified.
Phase II deliverables include detailed reports with relevant test data along with proof- of- concept hardware and components developed.

State of the Art and Critical Gaps

State of the Art photovoltaic array technology consists of high efficiency, multijunction cell technology on thick honeycomb panels. Lightweight arrays are just beginning to be developed. There are very limited demonstrated technology for High Intensity High Temperature (HIHT), Low Intensity Low Temperature (LILT), Solar Electric Propulsion (SEP) missions and Lunar orbital, surface or transfer applications.

Significant improvements in overall solar array performance are needed to address the current gaps between SOA (Sate of the Art) and many mission requirements for photovoltaic cell efficiency greater than 30%, array mass specific power greater than 300W/ kg, decreased stowed volume, reduced initial and recurring costs, long- term operation in radiation environments, high power arrays and a wide range of space, lunar, and planetary environmental operating conditions.

Relevance / Science Traceability

These technologies are relevant to any space science, earth science, planetary surface, or other science mission that requires affordable high-efficiency photovoltaic power production for orbiters, flyby craft, landers and rovers. Specific requirements can be found in the references listed above, but include many future Science Mission Directorate (SMD) missions. Specific requirements for orbiters and flybys to Outer planets include: LILT capability (>38% at 10 AU and <−140°C), radiation tolerance (6e15 1 MeV e-cm^2), high power (>50 kW at 1 AU), low mass (3× lower than SOP), low volume (3× lower than SOP), long life (>15 years), and high reliability.

These technologies are relevant and align to any Space Technology Mission Directorate (STMD) or Human Exploration and Operations Mission Directorate (HEOMD) mission that requires affordable high-efficiency photovoltaic power production.

NASA outlines New Lunar Science, Human Exploration Missions, found at:
https://www.nasa.gov/feature/nasa-outlines-new-lunar-science-human-exploration-missions

NASA Science Missions, found at:
https://science.nasa.gov/missions-page?field_division_tid=All&field_phase_tid=3951

NASA has plans to purchase services for delivery of payloads to the Moon through the Commercial Lunar Payload Services (CLPS) contract. Under this subtopic, proposals may include efforts to develop payloads for flight demonstration of relevant technologies in the lunar environment. The CLPS payload accommodations will vary depending on the particular service provider and mission characteristics. Additional information on the CLPS program and providers can be found at this link: https://www.nasa.gov/content/commercial-lunar-payload-services. CLPS missions will typically carry multiple payloads for multiple customers. Smaller, simpler, and more self-sufficient payloads are more easily accommodated and would be more likely to be considered for a NASA-sponsored flight opportunity. Commercial payload delivery services may begin as early as 2020 and flight opportunities are expected to continue well into the future. In future years it is expected that larger and more complex payloads will be accommodated. Selection for award under this solicitation will not guarantee selection for a lunar flight opportunity.

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Scope Description

NASA's Planetary Science Division is working to implement a balanced portfolio within the available budget and based on a decadal survey that will continue to make exciting scientific discoveries about our solar system. This balanced suite of missions shows the need for low mass/volume energy storage that can effectively operate in extreme environments for future NASA Science Missions.

Future science missions will require advanced primary and secondary battery systems capable of operating at temperature extremes from -200° C for outer planet missions to 400 to 500° C for Venus missions, and a span of -230° C to +120° C for missions to the Lunar surface. Operational durations of 60 days for Titan and 14 days for the Moon are of interest. Advancements to battery energy storage capabilities that address operation at extreme temperatures combined with high specific energy and energy density (>200 Wh/kg and >200 Wh/l) are of interest in this solicitation.

In addition to batteries, other advanced energy storage/load leveling technologies designed to the above mission requirements, such as mechanical or magnetic energy storage devices, are of interest. These technologies have the potential to minimize the size and mass of future power systems.

NASA has plans to purchase services for delivery of payloads to the Moon through the Commercial Lunar Payload Services (CLPS) contract. Under this subtopic, proposals may include efforts to develop payloads for flight demonstration of relevant technologies in the lunar environment. The CLPS payload accommodations will vary depending on the particular service provider and mission characteristics. Additional information on the CLPS program and providers can be found at this link: https://www.nasa.gov/content/commercial-lunar-payload-services. CLPS missions will typically carry multiple payloads for multiple customers. Smaller, simpler, and more self-sufficient payloads are more easily accommodated and would be more likely to be considered for a NASA-sponsored flight opportunity. Commercial payload delivery services may begin as early as 2020 and flight opportunities are expected to continue well into the future. In future years it is expected that larger and more complex payloads will be accommodated. Selection for award under this solicitation will not guarantee selection for a lunar flight opportunity.

References

• NASA Science:https://science.nasa.gov/
• Solar Electric Propulsion: https://www1.grc.nasa.gov/space/sep/

Expected TRL or TRL range at completion of the project: 3 to 5

Desired Deliverables of Phase II

Prototype

Desired Deliverables Description

Research should be conducted to demonstrate technical feasibility during Phase I and show a path toward a Phase II, and when possible, deliver a demonstration unit for NASA testing at the completion of the Phase II contract. Phase II emphasis should be placed on developing and demonstrating the technology under relevant test conditions. Additionally, a path should be outlined that shows how the technology could be commercialized or further developed into science-worthy systems.

State of the Art and Critical Gaps

State-of the-art primary and rechargeable cells are limited in both capacity and temperature range.  Typical primary Li-SO₂ and Li-SOCl₂ operate within a max temperature range of -40 to 80 deg C but suffer from capacity loss, especially at low temperatures.  At -40 deg C, the cells will provide roughly half the capacity available at room temperature. Similarly, rechargeable Li-ion cells operate within a narrow temperature range of -20 to 40 C and also suffer from capacity loss at lower temperatures. The lower limit of temperature range of rechargeable cells can be extended through the use of low temperature electrolytes, but with limited rate capability and concerns over lithium plating on charge. There is currently a gap that exists for high temperature batteries, primary and rechargeable, that can operate at Venus atmospheric temperatures. This solicitation is aimed at the development of cells that can maintain performance at extreme temperatures so as to minimize or eliminate the need for strict thermal management of the batteries, which adds complexity and mass to the spacecraft.

Relevance / Science Traceability

These batteries are applicable over a broad range of science missions.  Low temperature batteries are needed for potential NASA decadal missions to Ocean Worlds (Europa, Enceladus, and Titan) and the Icy Giants (Neptune, Uranus). These batteries are also needed for science missions on the lunar surface. Low temperature batteries developed under this subtopic would enhance these missions and could be potentially enabling if the missions are mass or volume limited.  There is also significant interest in a Venus surface mission that will require primary and/or rechargeable batteries that can operate for 60+ days on the surface of Venus. A high temperature battery that can meet these requirements is enabling for this class of missions.

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Scope Title

Innovative ways to transmit high power for lunar & Mars surface missions

Scope Description

The Global Exploration Roadmap (January 2018) and the Space Policy Directive (December 2017) detail NASA’s plans for future human-rated space missions. A major factor in this involves establishing bases on the lunar surface and eventually Mars. Surface power for bases is envisioned to be located remotely from the habitat modules and must be efficiently transferred over significant distances. The International Space Station (ISS) has the highest power (100kW), and largest space power distribution system with eight interleaved micro-grids providing power functions similar to a terrestrial power utility. Planetary bases will be similar to the ISS with expectations of multiple power sources, storage, science, and habitation modules, but at higher power levels and with longer distribution networks providing interconnection. In order to enable high power (>100kW) and longer distribution systems on the surface of the moon or Mars, NASA is in need of innovative technologies in the areas of lower mass/higher efficiency power electronic regulators, switchgear, cabling, connectors, wireless sensors, power beaming, power scavenging, and power management control. The technologies of interest would need to operate in extreme temperature environments, including lunar night, and could experience temperature changes from -153C to 123C for lunar applications, and -125C to 80C for Mars bases. In addition to temperature extremes, technologies would need to withstand (have minimal degradation from) lunar dust/regolith, Mars dust storms, and space radiation levels.

While this subtopic would directly address the lunar and Mars base initiatives, technologies developed could also benefit other NASA Mission Directorates including SMD (Science Mission Directorate) and ARMD (Aeronautics Research Mission Directorate). Specific projects which could find value in the technologies developed herein include Gateway, In-Situ Resource Utilization (ISRU), Advanced Modular Power Systems (AMPS), In-Space Electric Propulsion (ISP), planetary exploration, and Hybrid Gas Electric Propulsion. The power levels may be different, but the technology concepts could be similar, especially when dealing with temperature extremes and the need for electronics with higher power density and efficiency.

Specific technologies of interest would need to address the lunar or Mars environment, and include:

• Application of wide bandgap electronics in DC-DC isolating converters with wide temperature (-70ºC to 150ºC), high power density (>2 kW/kg), high efficiency (>96%) power electronics and associated drivers for voltage regulation.
• Low mass, highly conductive wires and terminations that provide reliable small gauges for long distance power transmission in the 1-10kW range, low mass insulation materials with increased dielectric breakdown strength and void reductions with 600 V or greater ratings, and low loss/low mass shielding.
• Power beaming concepts to enable highly efficient flexible/mobile power transfer in the 100-1,000W range, including the fusion of power/communication/navigation.

(See Z13.02 - Dust Tolerant Mechanismssubtopic to propose power connection/termination related technologies that are impervious to environmental dust and enable robotic deployment, such as robotically-enabled high voltage connectors and/or near-field wireless power transfer in the 1-10kW range.)

References

The Global Exploration Roadmap, January 2018: https://www.nasa.gov/sites/default/files/atoms/files/ger_2018_small_mobile.pdf

Space Policy Directive, December 2017: https://www.nasa.gov/topics/moon-to-mars/overview

Expected TRL or TRL range at completion of the project 3 to 6

Desired Deliverables of Phase II

Prototype, Analysis, Hardware, Research

Desired Deliverables Description

Typically, deliverables under Phase I proposals are geared towards a technology concept with associated analysis and design. A final report usually suffices in summarizing the work. Phase II hardware prototypes will have opportunities for infusion into NASA technology testbeds and commercial landers.

State of the Art and Critical Gaps

While high power terrestrial distribution systems exist, there is no equivalent to a lunar or planetary base. Unique challenges must be overcome in order to enable a realistic power architecture for these future applications, especially when dealing with the environmental extremes which will be encountered. The temperature swings will be a critical requirement on any technology developed, from power converters to cabling or power beaming concepts. In addition, proposals will have to consider lunar regolith and Mars dust storms.

Relevance / Science Traceability

This subtopic would directly address the lunar and Mars surface initiatives. There are potential infusion opportunities with SMD (Science Mission Directorate) Commercial Lander Payload Services and HEOMD (Human Exploration and Operations Mission Directorate) Flexible Lunar Exploration (FLEx) Landers. In addition, technologies developed could benefit other NASA missions including Gateway. The power levels may be different, but the technology concepts could be similar, especially when dealing with temperature extremes.

 

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Scope Title

Robotic Mobility, Manipulation and Sampling

Scope Description

Technologies for robotic mobility, manipulation, and sampling are needed to enable access to sites of interest and acquisition and handling of samples for in-situ analysis or return to Earth from planets and small planetary bodies. The Moon and planetary moons with liquid oceans are of particular interest, as well as Mars, comets, and asteroids.

Mobility technologies are needed to enable access to steep and rough terrain for planetary bodies where gravity dominates, such as Earth’s moon and Mars. Wheeled, legged, and aerial solutions are of interest. Wheel concepts with good tractive performance in loose sand while being robust to harsh rocky terrain are of interest. Technologies to enable mobility on small bodies and access to liquid below the surface (e.g., in conduits or deep oceans) are desired, as well as the associated sampling technologies. Manipulation technologies are needed to deploy sampling tools to the surface, transfer samples to in-situ instruments and sample storage containers, and hermetically seal sample chambers. Sample acquisition tools are needed to acquire samples on planetary and small bodies through soft and hard materials, including ice. Minimization of mass and ability to work reliably in the harsh mission environment are important characteristics for the tools. Finally, design for planetary protection and contamination control is important for sample acquisition and handling systems.

Component technologies for low-mass and low-power systems tolerant to the in-situ environment, e.g. temperature, radiation, and dust, are of particular interest. Technical feasibility should be demonstrated during Phase I and a full capability unit of at least TRL 4 should be delivered in Phase II. Proposals should show an understanding of relevant science needs and engineering constraints and present a feasible plan to fully develop a technology and infuse it into a NASA program. Specific areas of interest include the following:

• Surface mobility and sampling systems for planets, small bodies, and moons
• Near subsurface sampling tools such as icy surface drills to 30 cm depth deployed from a manipulator
• Subsurface ocean access such as via a deep drill system
• Sample handling technologies that minimize cross contamination and preserve mechanical integrity of samples
• Pneumatic sample transfer systems and particle flow measurement sensors
• Low mass/power vision systems and processing capabilities that enable fast surface traverse
• Active lighting stereo systems for landers and rovers
• Electro-mechanical connectors enabling tool change-out in dirty environments
• Tethers and tether play-out and retrieval systems
• Miniaturized flight motor controllers
• Cryogenic operation actuators
• Robotic arms for low gravity environments

Proposers should also note a related subtopic exists that is focused solely on lunar robotic missions (see Z5.05, Lunar Rover Technologies for In-Situ Resource Utilization and Exploration), under the Space Technology Mission Directorate). With NASA's present emphasis on lunar exploration, Z5.05 is provided to help develop innovative lunar rover technologies to support in-situ resource utilization activities and for developing ideas, subsystem components, software tools, and prototypes that contribute to more capable and/or lower cost lunar robots. In particular, cryogenic or cryo-capable actuators that are specifically for lunar rover applications should be directed towards Z5.05.

References

Mars Exploration/Programs & Missions: https://mars.nasa.gov/programmissions/

Solar System Exploration: https://solarsystem.nasa.gov/

Ocean Worlds website: https://www.nasa.gov/specials/ocean-worlds/

Ocean Worlds article: https://science.nasa.gov/news-articles/ocean-worlds

Expected TRL or TRL range at completion of the project: 2 to 4

Desired Deliverables of Phase II

Prototype, Analysis, Hardware, Software, Research

Desired Deliverables Description

Hardware and software for component robotic systems.

State of the Art and Critical Gaps

Scoops, powder drills, and rock core drills and their corresponding handling systems have been developed for sample acquisition on Mars and asteroids. Non-flight systems have been developed for sampling on comets, Venus, and Earth's moon. However, these have not been incorporated in a robotic mission, and the lack of a sufficient solution or technology readiness level is in some cases the reason a mission has not yet been possible. Exploration of icy ocean worlds is in the concept phase and associated sampling and sample handling systems do not exist.

Relevance / Science Traceability

The subtopic supports multiple programs within Science Mission Directorate (SMD). The Mars program has had infusion of technologies such as a force-torque sensor in the Mars 2020 mission. Recent awards would support the Ocean Worlds program: surface and deep drills for Europa. Products from this subtopic have been proposed for New Frontiers program missions. With renewed interest in return to Earth's moon, the mobility and sampling technologies could support future robotic missions to the moon.

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Scope Title

Improve the capability or performance of intravehicular activity robots

Scope Description

To support human exploration beyond Earth orbit, NASA is preparing to develop the "Gateway", which will be an orbiting facility near the Moon. This facility would serve as a starting point for missions to cis-lunar space and beyond. This facility could enable assembly and servicing of telescopes and deep-space exploration vehicles. This facility could also be used as a platform for astrophysics, Earth observation, heliophysics, and lunar science.

In contrast to the ISS (International Space Station), which is continuously manned, the Gateway is expected to only be intermittently occupied by humans – perhaps only 1 month per year. Consequently, there is a significant need for the Gateway to have autonomous capabilities for performing payload operations and spacecraft caretaking, particularly when astronauts are not present. Intra-Vehicular Activity (IVA) robots can potentially perform a wide variety of tasks including systems inspection, monitoring, diagnostics and repair, logistics and consumables stowage, exploration capability testing, aggregation of robotically returned destination surface samples, and science measurements and ops.

The objective of this subtopic, therefore, is to develop technologies that can improve the capability or performance of IVA robots to perform payload operations and spacecraft caretaking. Proposals are specifically sought to create technologies that can be integrated and tested with the NASA Astrobee or Robonaut 2 robots in the following areas: (1) Sensors and perception systems for interior environment monitoring, inspection, modeling and navigation; (2) Robotic tools for manipulating logistics and stowage or performing maintenance, housekeeping or emergency management operations (e.g. fire detection & suppression in multiple constrained locations or cleaning lunar dust out of HEPA (High-Efficiency Particulate Air) filters; and (3) Operational subsystems that enable extended robot operations (power systems, efficient propulsion, etc.), increase robot autonomy (planning, scheduling, and task execution), or improve human-robot teaming (software architecture, remote operations methods, etc.).

References

What is Astrobee? - https://www.nasa.gov/astrobee

What is a Robonaut? - https://www.nasa.gov/robonaut2

J. Crusan, et al. 2018. "Deep space gateway concept: Extending human presence into cislunar space", In Proceedings of IEEE Aerospace Conference, Big Sky, MT

M. Bualat, et al. 2018. "Astrobee: A new tool for ISS operations". In Proceedings of AIAA SpaceOps, Marseille, France.

T. Fong, et al. 2013. "Smart SPHERES: a telerobotic free-flyer for intravehicular activities in space". In Proceedings of AIAA Space 2013, San Diego, CA.

M. Diftler, et al. 2011. "Robonaut 2 - The first humanoid robot in space". In Proceedings of IEEE International Conference on Robotics and Automation, Shanghai, China.

M. Deans, et al. 2019. "Integrated System for Autonomous and Adaptive Caretaking (ISAAC)". Presentation, Gateway Intra-Vehicular Robotics Working Group Face to Face, Houston, TX; NASA Technical Reports Server [https://ntrs.nasa.gov/search.jsp?R=20190029054]

Expected TRL or TRL range at completion of the project: 4 to 5

Desired Deliverables of Phase II

Prototype, Analysis, Hardware, Software, Research

Desired Deliverables Description

Prototype components or subsystems. Proposals must describe how the technology will make a significant improvement over the current state of the art, rather than just an incremental enhancement, for a specific IVA robot application.

State of the Art and Critical Gaps

The technology developed by this subtopic would both enable and enhance the Astrobee free-flying robot and Robonaut 2 humanoid robot, which are the SOA for IVA robots. SBIR technology would improve the capability and performance of these robots to routinely and robustly perform IVA tasks, particularly internal spacecraft payload operations and logistics. New technology created by 2020 SBIR awards can be tested with these robots in ground testbeds at ARC and JSC during the SBIR period of performance. On-orbit testing on ISS may be possible during Phase 2 and beyond (Phase 2-E, 2-X, 3, etc.).

The technology developed by this subtopic would also fill technical gaps identified by the proposed GCD (Game Changing Development) "Integrated System for Autonomous and Adaptive Caretaking" (ISAAC) project, which will mature autonomy technology to support the caretaking of human exploration spacecraft. In particular, the SBIR technology would help provide autonomy and robotic capabilities that are required for in-flight maintenance (both preventive and corrective) of Gateway during extended periods when crew are not present.

Relevance / Science Traceability

This subtopic is directly relevant to the following STMD (Space Technology Mission Directorate) investments:

• Astrobee free-flying robot – GCD
• Integrated System for Autonomous and Adaptive Caretaking (ISAAC) – GCD
• Deep Space Smart Habitats – Space Technology Research Institutes (STRI)

This subtopic is directly relevant to the following HEOMD (Human Exploration and Operations Mission Directorate) investments:

• SPHERES/Astrobee facility – ISS
• Robonaut 2 humanoid robot – ISS
• Gateway program – Advanced Exploration Systems (AES)
• Logistics Reduction project – AES

Autonomous Systems Operations project – AES

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Scope Title

Carbon Dioxide Reduction System Components and Unit Processes

Scope Description

NASA has invested in many carbon dioxide reduction technologies over the years to increase the percentage of oxygen recovery from carbon dioxide in human spacecraft for long duration missions. Examples of technologies include, but are not limited to, Series-Bosch, Continuous Bosch, Methane Pyrolysis and Microfluidic Carbon Dioxide Electrolysis. Significant technical challenges still face these process technologies and are impeding progress in technology maturation. Critical technical elements of these technologies have a high degree of technical difficulty. Examples where additional technology development is needed include (this is a partial list):

• High temperature gas purification and/or separation for CO, H₂, and hydrocarbon rich streams.
• Nuisance particulate carbon contamination.
• Solid carbon clogging of frits and filters in recycle gas streams.
• Safe collection, removal and disposal of solid carbon while reactors are in operation.
• Subsystems to recharge reactors with new catalyst and to efficiently use or recycle consumable catalysts.

This subtopic is open to consider novel ideas that address any of the numerous technical challenges that face development of carbon dioxide reduction hardware with particular attention to those listed above. Specifics on two of these challenges are provided below.

Gas Purification and/or Separation for Carbon Monoxide, Hydrogen and Hydrocarbon Rich Streams

Many process technologies currently under development have challenging multi-component streams which could benefit from improved gas separation technology. High purity, high yield and continuous supply of separated gases are all desirable features of a proposed technology. The targeted process streams that may benefit from improved gas separations are the following:

• Producing a high-purity hydrogen product from a hydrogen-rich gas stream containing acetylene (as high as 6.4 mole %), trace amounts of other hydrocarbons (ethylene, ethane, benzene), unreacted methane, carbon monoxide, carbon dioxide and water vapor. It is imperative that the proposed separation technologies do not hydrogenate hydrocarbons, such as acetylene. This separation is directed at methane pyrolysis technologies including the Plasma Pyrolysis Assembly (PPA).
• Hydrogen separation from an ethylene-rich stream. This separation is directed at the effluent stream from a Microfluidic Electrochemical Reactor which consists of ethylene, hydrogen, methane, carbon monoxide, carbon dioxide and water vapor.
• Recovery of unreacted carbon dioxide and hydrogen from a carbon monoxide-rich stream. This separation is needed for a Bosch/Reverse Water Gas Shift (RWGS) Reactor.

Technology solutions could include, but not be limited to, filtration, mechanical separation or novel sorbents. If novel sorbents are developed the proposed technology solution should also address issues with scale-up to kg quantities (difficult for some novel sorbents). Technology solutions proposed in this subtopic could potentially be leveraged for In-Situ Resource Utilization (ISRU) applications.

Separation of Particulate Carbon and Hydrocarbons from Process Gas Streams

Oxygen recovery technology options, including carbon formation reactors and methane pyrolysis reactors almost universally result in particulates in the form of solid carbon or solid hydrocarbons. Mitigation for these particulates will be essential to the success and maintainability of these systems during long duration missions. Techniques and methods leading to compact, regenerable devices for removing, managing and disposing of residual particulate matter within ECLSS process equipment are sought. Separation performance approaching HEPA rating is desired for ultrafine particulate matter with minimal pressure drop. The separator should be capable of operating for hours at high particle loading rates and then employ techniques and methods to restore its capacity back to nearly 100% of its original clean state through in-place and autonomous regeneration or self-cleaning operations using minimal or no consumables (including media-free hydrodynamic separators). The device must minimize crew exposure to accumulated particulate matter and enable easy particulate matter disposal or chemical repurposing.

State of the Art and Critical Gaps

Future long duration human exploration missions may benefit from further closure of the Atmosphere Revitalization System (ARS).  The state-of-the-art Sabatier system, which has flown on the International Space Station as the Carbon Dioxide Reduction Assembly (CRA), only recovers about half of the oxygen from metabolic carbon dioxide. This is because there is insufficient hydrogen to react all available carbon dioxide. The Sabatier reacts hydrogen with carbon dioxide to produce methane and water.  The methane is vented overboard as a waste product causing a net loss of hydrogen. Mars missions target >75% oxygen recovery from carbon dioxide, with a goal to approach 100% recovery. NASA is developing several alternate technologies that have the potential to increase the percentage of oxygen recovery from carbon dioxide, toward fully closing the ARS loop. Methane pyrolysis recovers hydrogen from methane, making additional hydrogen available to react with carbon dioxide. Other technologies under investigation process carbon dioxide, recovering a higher percentage of oxygen than the Sabatier. All of these alternative systems, however, need additional technology investment to reach a level of maturity necessary for consideration for use in a flight environmental control and life support system (ECLSS).

 

Scope Title

Solid Carbon Repurposing

Scope Description

Solid carbon is produced as a major by-product from many candidate oxygen recovery technologies under consideration for long-duration missions, including Bosch, Series Bosch, Methane Pyrolysis by Carbon Vapor Deposition, and technologies containing carbon formation reactors. Based on metabolic CO₂ production for a crew of 4, 1.135 kg of solid carbon, with a volume as high as 2.8 liters, may be produced each day by oxygen recovery technologies, which then must be disposed of or repurposed. Repurposing of this carbon reduces logistical challenges associated with its disposal and may ultimately result in materials or processes advantageous for long-duration missions. The produced solid carbon may include nanofibers, microfibers and amorphous material with varying particle size, with the smallest in the micrometer range (10-50 µm). It may contain quantities of metals including, but not limited to, iron, nickel and cobalt. The solid carbon may be in the form of a loose powder or a densified cake with densities ranging from 0.4 to 1.8 g/cc and will vary by technology. Venting or disposal of this carbon to space will present considerable logistical challenges and will result in large volumes of space debris. Disposal of this carbon on a planetary surface may result in concerns for planetary protection or planetary science. NASA is seeking technologies and/or processes that repurpose solid carbon and its contaminants resulting in useful products for transit, deep space or planetary surface missions. The technology and/or process must limit crew exposure to the raw carbon.

References for All Scopes

"Hydrogen Recovery by Methane Pyrolysis to Elemental Carbon" (49th International Conference on Environmental Systems, ICES-2019-103)

"Evolving Maturation of the Series-Bosch System" (47th International Conference on Environmental Systems, ICES-2017-219)

"State of NASA Oxygen Recovery" (48th International Conference on Environmental Systems, ICES-2018-48)

"Particulate Filtration from Emissions of a Plasma Pyrolysis Assembly Reactor Using Regenerable Porous Metal Filters" (47th International Conference on Environmental Systems, ICES-2017-174)

"Methane Post-Processing and Hydrogen Separation for Spacecraft Oxygen Loop Closure" (47th International Conference on Environmental Systems, ICES-2017-182)

“Trading Advanced Oxygen Recovery Architectures and Technologies” (48th International Conference on Environmental Systems, ICES-2018-321)

NASA-STD-3001, VOLUME 2, REVISION A, Section 6.4.4.1 “For missions longer than 14 days, the system shall limit the concentration in the cabin atmosphere of particulate matter ranging from 0.5 μm to 10 μm (respirable fraction) in aerodynamic diameter to <1 mg/m3 and 10 μm to 100 μm to <3 mg/m3.”  https://www.nasa.gov/sites/default/files/atoms/files/nasa-std-3001-vol-2a.pdf.

Expected TRL or TRL range at completion of the project for Phase I: 3

Expected TRL or TRL range at completion of the project for Phase II for All Scopes:  4 to 5

Desired Deliverables of Phase II for All Scopes

Prototype, Analysis, Hardware, Research

Desired Deliverables Description for All Scopes

Phase I Deliverables - Reports demonstrating proof of concept, test data from proof of concept studies, concepts and designs for Phase II. Phase I tasks should answer critical questions focused on reducing development risk prior to entering Phase II. Conceptual solution in Phase I should look ahead to satisfying the requirement of limiting crew exposure to the raw carbon dust.

Phase II Deliverables - Delivery of technologically mature hardware, including components and subsystems that demonstrate performance over the range of expected spacecraft conditions. Hardware should be evaluated through parametric testing prior to shipment. Reports should include design drawings, safety evaluation, test data and analysis. Prototypes must be full scale unless physical verification in 1-g is not possible. Robustness must be demonstrated with long term operation and with periods of intermittent dormancy. System should incorporate safety margins and design features to provide safe operation upon delivery to a NASA facility.

State of the Art and Critical Gaps

No existing operational technology exists in this focused technical area. A crew of 6 during a 540 day Mars surface mission could potentially generate 920 kg of solid carbon - this will be a significant storage or disposal issue and may be a considerable raw product resource for potential utilization. Very limited research and development have been performed in this area. Some studies added carbon to plastic trash which subsequently was processed by a heat melt compactor to make "tiles", which encapsulated the carbon. Although these tiles are a safe way to get rid of trash waste, they were also studied for potential benefit for use as spacecraft radiation shielding.  Other work included adding binders to make rudimentary bricks for structural use.

Relevance / Science Traceability

These technologies would be essential and enabling to long duration human exploration missions, in cases where closure of the atmosphere revitalization loop will trade over alternate ECLSS architectures. The atmosphere revitalization loop on the ISS is only about 50% closed when the Sabatier is operational. These technologies may be applicable to Gateway, Lunar surface, and Mars, including surface and transit. This technology could be proven on the ISS.

This subtopic is directed at needs identified by the Life Support Systems Capability Leadership Team (CLT) in areas of water recovery and environmental monitoring, functional areas of Environmental Control and Life Support Systems (ECLSS).

The Life Support Systems (LSS) Project, under the Advanced Exploration Systems (AES) Program, within the Human Exploration and Operations Mission Directorate (HEOMD), is the expected customer. The LSS Project would be in position to sponsor Phase III and technology infusion.

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As the design for the new Exploration Extra-vehicular Mobility Unit (xEMU) is developed, there are obvious gaps in technologies, which need to be fulfilled to meet the new exploration requirements. Various Exploration Portable Life Support System (xPLSS) Hatch components are at a stall in technology development and require new innovative ideas. These xPLSS Hatch Components (through three scopes) are the focus areas for this solicitation in an attempt to integrate new technologies into the xPLSS.  NASA has plans to go to the moon and as the mission extends further out of Lower Earth Orbit, durability and extensibility will become some of the most important requirements.

This subtopic is relevant to the Exploration Extravehicular Mobility Unit (xEMU), ISS, as well as commercial space companies. As a new Space Suit Exploration Portable Life Support System (xPLSS) is being designed, built, integrated and tested at JSC and integrated into the xEMU, solutions will have a direct infusion path as the xPLSS is matured to meet the design and performance goals.

Scope Title

Feedwater Supply Assembly

Scope Description

Sterile compliant bladder, capable of storing ultrapure feedwater with a relatively high cycle life:In order for the thermal control loop to operate properly, a water source is needed. An effective, efficient, sterile and durable feedwater bladder is essential. The suit pressure acts on this bladder and as water evaporates, the bladder resupplies the loop. The bladder must be clean and not leak particulates or polymer chains over long periods of quiescence. The water in the control loop contains a biocide and the bladder must not react with these chemicals to form potential contaminants. The maximum design pressure (MDP) for the system at a lunar environment will be 16 psid with a cycle life of 4 X 156 = 624 MDP. Having a bladder with these qualities not only buys down the safety risk of rupture, it promotes reliability at higher pressures and provides an avenue to extend Extravehicular Activity (EVA) length.

References

Feedwater Supply Assembly Requirements

Note to vendor: The following two drawings referenced in the above specification shall be provided if vendor is selected for award.

1. Feedwater Supply Assembly (FSA 431) Drawing SLN 13102397 
   

   https://ntrs.nasa.gov/search.jsp?R=20190033446

2. Auxiliary Feedwater Supply Assembly (FSA 531) Drawing SLN 13102398 
   

   https://ntrs.nasa.gov/search.jsp?R=20190033446

 

Scope Title

Bypass Relief Valve

Scope Description

Material dependent Relief Valve (RV) capable of re-calibration: The bypass relief valve cracks and flows from the pump outlet to the pump inlet, short-circuiting the pump when there is a blockage in the line. It is a safety feature designed to limit the head pressure that could be generated by the positive displacement pump, which is used in the primary and auxiliary thermal control loops. Materials, design pressures and re-calibration capabilities are a priority for this design. The desired housing material is titanium, which is a difficult metal to work with, but is a requirement as a preventative measure to avoid galvanic coupling between interfacing metals. To ensure the thermal loop pressure stays within a safe range, the crack and reseat pressures must be between 14-15 psid with a full flow of 220 lb/hr at <18 psid. The design should also include a method of setting or re-calibrating the cracking pressure in case there is drift over time. Replacement of the entire unit is not preferred due to accessibility and operational concerns.

References

Thermal Loop Bypass Relief Valve Requirements

Note to vendor: The following drawing referenced in the above specification shall be provided if vendor is selected for award.

• Bypass Relief Valve Assembly (RV-424/RV-524) SLN13102925
  

  https://ntrs.nasa.gov/search.jsp?R=20190033446

 

Scope Title

Trace Contaminant Control

Scope Description

Trace contaminant removal capability:  Non-regenerable activated carbon is the current state of the art for trace contamination control.  However, this provides a logistics impact to future missions. The primary trace contaminants that must be removed include ammonia (NH₃), carbon monoxide (CO), formaldehyde (CH₂O), and methanethiol (also known as methyl mercaptan) (CH₃SH). The minimum objective would be to remove all of the significant compounds that threaten to exceed the 7-day Spacecraft Maximum Allowable Concentrations (SMAC) values during an EVA. The ideal solution would be a vacuum-regenerable sorbent that could be integrated with the Exploration Portable Life Support System (xPLSS) CO₂/H₂O removal system. This system performs regeneration or desorption by exposing the sorbent to a pressure swing from 4.3 psia to <1 torr over approximately 2 minutes. Temperatures remain in the 60-80^oF range with a small amount of heat flux from the cross-coupled adsorbing bed. Additional heat input requirements from resistance heaters or other sources would negatively impact the system trade the more significant the value becomes.

References

Trace Contamination Control Cartridge Requirements

Note to vendor: The following drawing referenced in the above specification shall be provided if vendor is selected for award.

• Trace Contamination Control (TCC-360) Specification Control Drawing SLN13102266 
  

  https://ntrs.nasa.gov/search.jsp?R=20190033446

Expected TRL or TRL range at completion of the project for all scopes: 3 to 5

Desired Deliverables of Phase II for all scopes

Prototype

Desired Deliverables Description for all scopes

Phase I products:  By the end of Phase I, it would be beneficial to have a concept design for infusion into the Exploration Portable Life Support System (xPLSS). Testing of the concept is desired at this Phase.

Phase II products:  By the end of Phase II, a prototype ready for system-level testing in the xPLSS or in a representative loop of the PLSS is desired.

State of the Art and Critical Gaps

As the design for the new Exploration Extra-vehicular Mobility Unit (xEMU) is developed, there are obvious gaps in technologies, which need to be fulfilled to meet the new exploration requirements. Various Exploration Portable Life Support System (xPLSS) Hatch components are at a stall in technology development and require new innovative ideas. These xPLSS Hatch Components are the focus areas for this solicitation in an attempt to integrate new technologies into the xPLSS. NASA has plans to go to the moon and as the mission extends further out of Lower Earth Orbit, durability and extensibility will become some of the most important requirements.

Relevance / Science Traceability

It is relevant to the Exploration Extravehicular Mobility Unit (xEMU), ISS, as well as commercial space companies. As a new Space Suit Exploration Portable Life Support System (xPLSS) is being designed, built, integrated, and testing at JSC and integrated into the xEMU, solutions will have a direct infusion path as the xPLSS is matured in to meet the design and performance goals.

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Scope Title

Model Based Systems Engineering for Distributed Development

Scope Description

Systems Engineering technology is both a critical capability and a bottleneck for NASA human exploration development. NASA looks to a sustainable return to the Moon to enable future exploration of Mars, components such as Lunar Gateway and Commercial Lunar Payload Services (CLPS) will require partnerships with a wide variety of communities. Building from the success of the international partnerships for International Space Station (ISS), space agencies from multiple governments are looking for roles on the Gateway. A particular focus has been made to include the rapidly growing commercial space industry to provide an important role in supporting a sustained presence on the Moon. All of these potential partners will have their own design capabilities, their own development processes and internal constituencies to support. Integrating and enabling disparate systems built in different locations by different owners to all work cohesively together will require a significant upgrade to the core systems engineering capabilities.

In the last decade Model-Based Systems Engineering (MBSE) technology has matured as evidenced by the development of Systems Modeling Language (SysML) tools and frameworks that support engineers in development efforts from requirements through hardware and software implementation. MBSE holds considerable promise for accelerating, reducing overhead labor, and improving the quality of systems development. However, a remaining bottleneck is the coordination and integration of system development across distributed organizations, such as the multiple partners developing lunar gateway and eventual Mars exploration. This subtopic seeks technology to fill this gap.

Areas of particular need include:

• Methodologies that support integration among tools and exchange of information between multidisciplinary artifacts using automated intelligent reasoning.
• The definition of open interface standards and tools to enable inspection of distributed models across engineering domains.
• Tools or systems that allow models to be shared across development environments and trace the resulting system model back to contributions from multiple partners.
• Modeling environments that facilitate user interaction from multiple stakeholders of varying expertise in MBSE.
• Continuous integration and verification of safety critical system requirements that depend on disparate development sources.

References:

• https://www.nasa.gov/consortium/ModelBasedSystems
• http://www.omgsysml.org(link is external)
• Ensuring information exchange of digital artifacts are transferable and up to date among multiple stakeholders.
  • Digital Engineering Information Exchange Working Group (DEIX WG): http://www.omgwiki.org/MBSE/doku.php?id=mbse:deix(link is external)
• Computational tools to augment human decision making and reasoning on complex systems with large amounts of data from disparate sources
  • Augmented Intelligence for Systems Engineering challenge team (AI-SECT): http://www.omgwiki.org/MBSE/doku.php?id=mbse:augmented(link is external)
• Automated formal specification, formal verification, and test case generation of requirements with linked data and traceability to discipline specific (CAD, CAE, etc.) tools, particularly requirements with safety properties.
  • ReqIF: https://www.omg.org/reqif/(link is external)
  • SysPhs: https://www.omg.org/spec/SysPhS/(link is external)
  • FMI: https://fmi-standard.org(link is external) 
• Lightweight and intuitive cloud-based interfaces for CRUD (create, read, update, delete) operations on models particularly for users with limited MBSE experience.
• Open-MBEE: https://openmbee.org(link is external)
• OSLC: https://open-services.net/(link is external)

Expected TRL or TRL range at completion of the project: 4 to 6

Desired Deliverables of Phase II

Prototype, Software

Desired Deliverables Description

Methodologies and tools that support distributed development efforts

State of the Art and Critical Gaps

For distributed development, the state-of-the-art tends to be laboriously negotiated interface control documents and manual integration processes that are inherently slow and labor intensive. In an effort to overcome these challenges MBSE and SysML in particular has seen significant adoption at NASA (Gateway, Resource Prospector, Europa Clipper, Space Communications and Navigation [SCaN], Space Launch System [SLS]) especially after the MBSE Pathfinder ('16/'17) and MBSE Infusion And Modernization Initiative (MIAMI, '18/'19) studies. However, these pilot programs and a survey of NASA's use of MBSE conducted by NASA Independent Verification & Validation (IV&V) and Ames Research Center identified areas of critical need, including:

1. Sharing and version control of models.
2. Integration of SysML of domain specific tools
3. Steep learning curve for users with limited MBSE experience
4. Testing, Verification and Validation with SysML have limited use
5. No tools exist for formally specifying requirements and linking to model properties

With programs such as Gateway and Artemis that require coordination among multiple NASA centers, international space agencies, and commercial partnerships these needs will be amplified.   Tool infrastructures that enable integrated support of requirements tracing, design reference points, intelligent reasoning of data and interface constructs are generally not available except within proprietary boundaries. We need tools that support integrated development and model sharing across development environments and that support use across multiple vendors.

Relevance / Science Traceability

This subtopic would be of relevance to all Human Exploration and Operations Mission Directorate (HEOMD) missions, but of particular interest will be Gateway and Artemis development. Those systems have already adopted the use of MBSE tools and tools sought help reduce potential system integration bottlenecks. Over the next 3 to 5 years, there will be considerable opportunity for small business contributions to be matured and integrated into the support infrastructure as Gateway evolves from concept to development program.

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Scope Title

Nanotechnology Innovations for Spacecraft Water Management Applications

Scope Description

Water recovery from wastewater sources is key to long duration human exploration missions. Without substantial water recovery, life support system launch weights are prohibitively large. Regenerative systems are utilized on the International Space Station (ISS) to recycle water from humidity condensate and urine. The Water Processor Assembly (WPA) accepts distillate from the Urine Processor Assembly (UPA) and humidity condensate from condensing heat exchanges. The WPA contains multi-filtration beds to remove inorganic and non-volatile organic contaminants, followed by a catalytic oxidation reactor where low molecular weight organics not removed by the adsorption process are oxidized in the presence of oxygen, elevated temperature, and a catalyst. To stabilize urine and protect components from biofouling and precipitation, a toxic pretreatment formula is added to collected urine. Simple measurements of water composition are made during flight, including conductivity, total organic carbon and iodine concentration. For determination of ionic or organic species in water and wastewater, samples must be returned to earth.

This subtopic solicits for technologies to fill specific gaps in NASA’s water management systems for human spaceflight. Proposals must address needs in one of the three target areas specified. These areas of scope are aligned with the three specific thrusts described within the white paper of the Nanotechnology Signature Initiative (NSI) "Water Sustainability through Nanotechnology". Please see references for additional information, including water quality requirements and guidelines.

Increasing Water Availability Using Nanotechnology:  Removal of Problematic Contaminants from Processed Wastewater

Two problematic organic compounds are recalcitrant to WPA processing on the ISS. Dimethylsilanediol (DMSD) is a silicon-containing degradation byproduct from siloxane based compounds. DMSD can violate ISS potable water quality standards over time, requiring premature multifiltration (MF) bed replacement. Dimethyl sulfone (DMSO2) is a sulfur-containing metabolic byproduct that has historically been consistently present in ISS potable water delivered to the Oxygen Generation Assembly (OGA) for electrolysis to O₂ and H₂. DMSO2 accumulates in the OGA water recirculation loop and is thus present in the OGA hydrogen product stream. When fed to the Sabatier reactor this contaminated H₂ has been shown to poison the Sabatier catalyst over time from sulfur exposure. The presence of DMSO2 is negatively impacting exploration design requirements and Concepts of Operation (CONOPS) for the Advanced-OGA and the Sabatier subsystems, including periodic automated flushing and trace contaminant getter devices. The development of a technology or method for physicochemical removal of these contaminants, compatible with the ISS WRS/WPA, will benefit both current manned and future exploration missions. Although technical solutions are sought that involve novel utilization of nanotechnology, proposals using more conventional or alternative approaches will also be considered.

Improving the Efficiency of Water Delivery and Use with Nanotechnology:  Management and Monitoring of Silver Biocide in Potable Water

NASA is considering using silver as the active biocide in potable water systems for use in future spacecraft. NASA is seeking technologies for delivery, maintenance and monitoring silver in potable water.

• NASA seeks technologies to deliver and replenish silver ions in potable water, to maintain a concentration at a chosen set point within a range of 200 to 400 ug/L. The system should be capable of operating in-line, to deliver silver at a flow rate of 0.1 to 0.15 L/min potable water. Furthermore, the device should be able to operate at ambient temperature, pH ranges between 4.5 - 9.0, and system pressures up to 30 psig (200 kPa). Moreover, the device should also be small, robust, lightweight, and have minimal power and consumable mass requirements. Additionally, candidate technologies should be microgravity compatible and have no adverse effects on the potability of the drinking water system. The technology should also be capable of providing continuous, stable and autonomous operation, and be fully functional following periods of long-term system dormancy – up to 1 year.
• Silver ions may drop out of solution, depositing on fluid lines and tank surfaces, resulting in loss of silver concentration, impacting its efficacy as a residual disinfectant in potable water. Alternative methods are sought to minimize loss of silver ions in spacecraft potable water plumbing systems.
• NASA is interested in sensing technologies for the in-line measurement of ionic silver in spacecraft potable water systems. Overall, the sensing technology should offer small, robust, lightweight, low-power, compatible design solutions capable of stable, continuous, and autonomous measurements of silver for extended periods of time. Sensors of particular interest would provide:  continuous in-line measurement of ionic silver at concentrations between 0 and, at least, 1000 parts per billion (ppb); a minimum detection limit of 10 ppb or less; measurement accuracy of at least 2.5% full scale (1000 ppb); stable measurements in flows up to 0.5 L/min and pipe diameters up to ¾ inch; high sampling frequency, e.g., up to 1 measurement per minute; stable calibration, greater than 3 years preferred; minimal and/or no maintenance requirements; operation at ambient temperature, system pressures up to 30 psig (200 kPa), and a solution pH between 4.5 - 9.0; and finally, a volumetric footprint less than 2000 cubic centimeters. The sensing technology should have little to no impact on the overall volume and concentration of silver being maintained within the spacecraft water system.

Enabling Next-Generation Water Monitoring Systems with Nanotechnology

NASA is seeking miniature analytical systems to measure mineral and organic constituents in potable water and wastewater. NASA is interested in sensor suites capable of simultaneous measurement of inorganic and organic species. Spacecraft applications exist for monitoring species within wastewater (potential waste streams:  urine, humidity condensate, Sabatier product water, waste hygiene, and waste laundry water), regenerated potable water and in support of on-board science. Multi-species analyte measurement capability is of interest that would be competitive to standard water monitoring instruments such as ion-chromatography, inductively coupled plasma spectroscopy, and high performance liquid chromatography. Components that enable the miniaturization of these monitoring systems, such as microfluidics and small scale detectors, will be considered. Technologies should be targeted to have >3 year service life and >50% size reduction compared to current state of the art. Ideally, monitoring systems should require no hazardous reagents, have long-term calibration stability, and require very little crew time to operate and maintain.

References

NASA is a collaborating agency with the NTSC Committee on Technology Subcommittee on Nanoscale Science, Engineering and Technology's Nanotechnology Signature Initiative (NSI): "Water Sustainability through Nanotechnology" (Water NSI). For a white paper on the NSI, see https://www.nano.gov/node/1580(link is external)

A high-level overview of NASA's spacecraft water management was presented at a webinar sponsored by the Water NSI: "Water Sustainability through Nanotechnology: A Federal Perspective, Oct. 19, 2016" https://www.nano.gov/publicwebinars(link is external)

A general overview of the state of the art of spacecraft water monitoring and technology needs was presented at a webinar sponsored by the Water NSI: "Water Sustainability through Nanotechnology: Enabling Next-Generation Water Monitoring Systems, Jan. 18, 2017" located at https://www.nano.gov/publicwebinars(link is external)

For a list of targeted contaminants and constituents for water monitoring, see "Spacecraft Water Exposure Guidelines for Selected Waterborne Contaminants" located at https://www.nasa.gov/feature/exposure-guidelines-smacs-swegs

Advanced Exploration Systems Program, Life Support Systems Project https://www.nasa.gov/content/life-support-systems

National Aeronautics and Space Administration, NASA Technology Roadmaps, TA 6: Human Health, Life Support, and Habitation Systems (National Aeronautics and Space Administration, Draft, May 2015, www.nasa.gov/sites/default/files/atoms/files/2015_nasa_technology_roadmaps_ta_6_human_health_ life_support_habitation.pdf).

Layne Carter, Jill Williamson, Daniel Gazda, Chris Brown, Ryan Schaezler, Frank Thomas, Jesse Bazley, Sunday Molina “Status of ISS Water Management and Recovery” 49th International Conference on Environmental Systems, ICES-2019-36 https://ttu-ir.tdl.org/bitstream/handle/2346/84720/ICES-2019-36.pdf(link is external)

Dean L. Muirhead, Layne Carter “Dimethylsilanediol (DMSD) Source Assessment and Mitigation on ISS: Estimated Contributions from Personal Hygiene Products Containing Volatile Methyl Siloxanes (VMS)” 48th International Conference on Environmental Systems, ICES-2018-123. https://ttu-ir.tdl.org/bitstream/handle/2346/74112/ICES_2018_123.pdf(link is external)

Chad Morrison, Christopher McPhail, Mike Callahan, Stuart Pensinger “Concepts for a Total Organic Carbon Analyzer for Exploration Missions” 49th International Conference on Environmental Systems, ICES-2018-254 https://ttu-ir.tdl.org/bitstream/handle/2346/84465/ICES-2019-254.pdf(link is external)

Molly S. Anderson, Ariel V. Macatangay, Melissa K. McKinley, Miriam J. Sargusingh, Laura A. Shaw, Jay L. Perry, Walter F. Schneider, Nikzad Toomarian, Robyn L. Gatens " NASA Environmental Control and Life Support Technology Development and Maturation for Exploration: 2018 to 2019 Overview", 49th International Conference on Environmental Systems, ICES-2019-297 https://ttu-ir.tdl.org/bitstream/handle/2346/84496/ICES-2019-297.pdf(link is external)

Donald Layne Carter, David Tabb, Molly Anderson "Water Recovery System Architecture and Operational Concepts to Accommodate Dormancy", 47th International Conference on Environmental Systems, Paper ICES-2017-43 https://ttu-ir.tdl.org/ttu-ir/bitstream/handle/2346/72884/ICES_2017_43.pdf(link is external)

Li, Wenyan, Calle, Luz, Hanford, Anthony, Stambaugh, Imelda and Callahan, Michael "Investigation of Silver Biocide as a Disinfection Technology for Spacecraft – An Early Literature Review", 48th International Conference on Environmental Systems, Paper ICES-2018-82

Expected TRL or TRL range at completion of the project for Phase I:  3

Expected TRL or TRL range at completion of the project for Phase II:  4 to 5

Desired Deliverables of Phase II:

Research, Analysis, Prototype, Hardware

Desired Deliverables Description

Phase I Deliverables - Reports demonstrating proof of concept, including test data from proof of concept studies, and concepts and designs for Phase II. Phase I tasks should answer critical questions focused on reducing development risk prior to entering Phase II.

Phase II Deliverables - Delivery of technologically mature hardware, including components and subsystems that demonstrate performance over the range of expected spacecraft conditions. Hardware should be evaluated through parametric testing prior to shipment. Reports should include design drawings, safety evaluation, test data and analysis. Prototypes must be full scale unless physical verification in 1-g is not possible. Robustness must be demonstrated with long term operation and with periods of intermittent dormancy. System should incorporate safety margins and design features to provide safe operation upon delivery to a NASA facility.

State of the Art and Critical Gaps

NASA has unique water needs in space that have analogous applications on Earth. NASA’s wastewater collection differs from systems used on Earth in that it is highly concentrated with respect to urine, uses minimal flush water, is separated from solid wastes, and contains highly acidic and toxic pretreatment chemicals. NASA is interested in recovery of potable water from waste water, low toxicity residual disinfection, antifouling treatments for plumbing lines and tanks, "microbial check valves" that prevent microbial cross-contamination where water treatment and potable water systems share connections, and miniaturized sensors and monitoring systems for contaminants in potable water and waste water. NASA’s goal is zero-discharge water treatment, targeting 100% water recycling and reuse. Spacecraft traveling away from Earth require the capability of a fully functional water analysis laboratory, including identification and quantification of known and unknown inorganic ions, organics, and microbes, as well as pH, conductivity, total organic carbon and other typical measurements. Spacecraft Water Exposure Guidelines (SWEGs) have been published for selected contaminants. Nanotechnology may offer solutions in all of these application areas.

Relevance / Science Traceability

This technology could be proven on the ISS and would be useful to long duration human exploration missions, including Gateway, Lunar surface, and Mars, including surface and transit.  It is essential and enabling for water to be recycled to reduce launch costs associated with life support consumables. This subtopic is directed at needs identified by the Life Support Systems Capability Leadership Team (CLT) in areas of water recovery and environmental monitoring, functional areas of Environmental Control and Life Support Systems (ECLSS).

This subtopic is directed at meeting NASA's commitments as a collaborating agency in the National Nanotechnology Signature Initiative: "Water Sustainability through Nanotechnology". This initiative was established under the NTSC Committee on Technology, Subcommittee on Nanoscale Science, Engineering and Technology.

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Scope Description

This subtopic seeks technologies related to cryogenic propellant (e.g. hydrogen, oxygen, methane) production, storage, transfer, and usage to support NASA's in-situ resource utilization (ISRU) goals. This includes a wide range of applications, scales, and environments consistent with future NASA missions to the Moon and Mars. Anticipated outcome of Phase I proposals are expected to deliver proof of the proposed concept with some sort of basic testing or physical demonstration. Proposals shall include plans for a prototype and demonstration in a defined relevant environment (with relevant fluids) at the conclusion of Phase II. Solicited topics are as follows:

• Subgrid Computational Fluid Dynamics (CFD) model that would model spray transport heat transfer and wall interactions during spray heat transfer during cryogenic propellant tank chilldown and fill in microgravity. Three submodels should be developed, including a (1) droplet transport and heat and mass transfer model, (2) fluid-to-wall boiling model covering all pertinent regimes (flash evaporation, film boiling, transition boiling, nucleate boiling, condensation), and (3) model that is used to capture bulk phases (e.g., volume of fluid). There should be seamless coupling between all three submodels. Emphasis should be on cryogenic fluids such as liquid hydrogen, oxygen, methane, and nitrogen. Phase I should have an emphasis on 1-g while Phase II should include microgravity applications. Models must be anchored to experimental cryogenic data.
• Develop and demonstrate methodologies for recovering propellant from lunar and Martian descent stages that have low fill levels (< 5%) of liquid oxygen, hydrogen, and/or methane mixed with helium. Methodologies can assume liquid extraction (for a short amount of time) or vapor extraction. Possible uses of the fluids could include fuel cells, life support/breathing air, or other applications. Methodologies should focus on the amount of propellant that might be extractable at different purities (prop/helium). Phase I should focus on defining and refining the methodologies for scavenging, as well as defining what should be done to the landers to enable or facilitate later access for scavenging. Phase II should include some sort of a demonstration, perhaps using simulant or similar fluids.
• Develop and defend a proposed relaxed propellant grade specification for liquid oxygen, liquid methane, and/or liquid hydrogen, allowing higher amounts of water contaminants in the oxygen and hydrogen, and higher amounts of water, hydrogen, and carbon monoxide/dioxide in the methane. Starting with assessment of potential impurities coming out of the ISRU production plant, analysis should evaluate the effects on the liquefaction system, pump and pressure-fed propellant feed system, and engine performance, especially potential stability effects. Phase I should conclude with a proposed relaxed propellant specification for at least one propellant (oxygen or methane priority over hydrogen), with identification of the propulsion component (liquefaction, feed system, injectors, etc.) that has the most sensitivity to the impurities and will therefore drive the limits on the specification. Phase II should include a hardware demonstration of the critical element at a minimum to validate the accuracy of the analytical predictions.
• Advance non-liquid electrolyte technologies for chemical flow cells (e.g., fuel cells, electrolyzers, flow batteries, etc.) that generate electrical power from a chemical reaction or reconstitute a reaction byproduct into fuels and oxidizer for such a chemical flow cell. These electrolytes are required to be cycled through very low temperatures (< 150 K) during storage to survive a lunar night or cis-lunar travel and recover completely (>98%) mechanical, electrical, and chemical performance. Ideally, these electrolytes would be able to process propellants (hydrogen, oxygen, methane, kerosene, etc.) and either tolerate or recover from exposure to standard propellant contaminants with minimal/no performance loss. Due to the potential for high fluid pressures and vibration loads, any proposal will illustrate how the electrolyte could be mechanically supported to operate hermetically under these conditions. To demonstrate that the electrolyte exceeds the State of Art, the deliverable test article will support an electrical current density of at least 300 mA/cm2 for at least 500 hours, support transient currents > 750 mA/cm2 for at least 30 seconds, and support slew rates > 50 A/cm2/s. Providing test data for the electrolyte performance degradation rate when operated as intended is required with test times >5,000 hours significantly strengthening the proposal. It would be beneficial if the electrolyte operated reversibly with equal efficiently. Liquid electrolytes, loose or contained within a support structure, are excluded from this Scope due to the complications that liquid electrolytes pose for an eventual system during launch.

NASA has plans to purchase services for delivery of payloads to the Moon through the Commercial Lunar Payload Services (CLPS) contract. Under this subtopic, proposals may include efforts to develop payloads for flight demonstration of relevant technologies in the lunar environment. The CLPS payload accommodations will vary depending on the particular service provider and mission characteristics. Additional information on the CLPS program and providers can be found at this link: https://www.nasa.gov/content/commercial-lunar-payload-services. CLPS missions will typically carry multiple payloads for multiple customers. Smaller, simpler, and more self-sufficient payloads are more easily accommodated and would be more likely to be considered for a NASA-sponsored flight opportunity. Commercial payload delivery services may begin as early as 2020 and flight opportunities are expected to continue well into the future. In future years it is expected that larger and more complex payloads will be accommodated. Selection for award under this solicitation will not guarantee selection for a lunar flight opportunity.

References

1. Kartuzova, O., and Kassemi, M., "Modeling K-Site LH2 Tank Chilldown and no Vent Fill in Normal Gravity" AIAA-2017-4662

2. Chato, D. "LOX Tank Helium Removal for Propellant Scavenging Test" presentation at 2008 AIAA Aerospace Sciences Meeting, Orlando, FL, 2008.

3. Regenerative Fuel Cell Power Systems for Lunar and Martian Surface Exploration (https://arc.aiaa.org/doi/abs/10.2514/6.2017-5368(link is external))

4. NASA Technology roadmap (https://gameon.nasa.gov/about/space-technology-roadmap/), §TA03.2.2.1.2. Chemical Power Generation and §TA03.2.2.2.3. Regenerative Fuel Cell Energy Storage (NOTE: This may be a dated link as this Roadmap still references ETDP/ETDD)

5. Commercial Lunar Propellant Architecture: A Collaborative Study of Lunar Propellant Production (https://doi.org/10.1016/j.reach.2019.100026(link is external))

6. Linne, et.al. “Feasibility of Scavenging Propellants from Lander Descent Stage to Supply Fuel Cells and Life Support,” AIAA-2009-6511, September, 2009.    

Expected TRL or TRL range at completion of the project: 2 to 4

Desired Deliverables of Phase II

Prototype, Hardware, Software

Desired Deliverables Description

Phase I proposals should at minimum deliver proof of the concept, including some sort of testing or physical demonstration, not just a paper study. Phase II proposals should provide component validation in a laboratory environment preferably with hardware (or model subroutines) deliverable to NASA.

Electrolyte technologies for chemical cell product deliverables would be an operational electrochemical test article demonstrating the capability of the electrolyte to support the listed current density by processing the intended propellants when packaged as a flow cell. This test article will have an active area of at least 50 cm2 and would ideally contain multiple cells to demonstrate extensibility to existing stack designs. It would be favorable to include empirical electrochemical performance data of the electrolyte over as much of the pressure range from 5 psia to 3015 psia as possible to illustrate the potential viability range for Lunar applications.

State of the Art and Critical Gaps

Cryogenic Fluid Management is a cross-cutting technology suite that supports multiple forms of propulsion systems (nuclear and chemical), including storage, transfer, and gauging, as well as liquefaction of ISRU produced propellants.  Space Technology Mission Directorate (STMD) has identified that Cryogenic Fluid Management (CFM) technologies are vital to NASA's exploration plans for multiple architectures, whether it is hydrogen/oxygen or methane/oxygen systems including chemical propulsion and nuclear thermal propulsion. For spray transport and film condensation, there are significant gaps in modeling. For scavenging, only small scale tests have been conducted to remove residual helium from a liquid oxygen tank.

There is currently no standard on propellant grade specification for an ISRU plant.

Existing electrolytes for space applications are limited to a polymeric membrane based on perfluorinated teflon and ceramic electrolyte. While it has the necessary electrochemical and mechanical properties, the polymeric membrane has very tight thermal constraints due to a high moisture content which complicates thermal system designs for lunar systems during transit. It is also very sensitive to chemical contamination. The ceramic electrolyte has significant mechanical and slew rate limitations, but is more resilient to chemical contamination and has a much larger thermal range which allows storage in very cold environments. Once operational and at temperature, either existing electrolyte technology operates in cold lunar regions. Should an off-nominal event occur during the lunar night that results in a cold-soak, neither existing electrolyte technology has a meaningful chance of recovering from the exposure to the low temperatures.

Relevance / Science Traceability

STMD strives to provide the technologies that are needed to enable exploration of the solar system, both manned and unmanned systems; cryogenic fluid management is a key technology to enable exploration. Whether liquid oxygen/liquid hydrogen or liquid oxygen/liquid methane is chosen by Human Exploration and Operations Mission Directorate (HEOMD) as the main in-space propulsion element to transport humans, CFM will be required to store propellant for up to 5 years in various orbital environments. Transfer will also be required, whether to engines or other tanks (e.g., depot/aggregation), to enable the use of cryogenic propellants that have been stored. In conjunction with ISRU, cryogens will have to be produced, liquefied, and stored, the latter two of which are CFM functions for the surface of the Moon or Mars. ISRU and CFM liquefaction drastically reduces the amount of mass that has to be landed on the Moon or Mars.

NASA already has proton exchange-membrane (PEM) based electrochemical hardware in the International Space Station (ISS) Oxygen Generator Assembly and is developing electrochemical systems for space applications through the Evolved Regenerative Fuel Cell. These system designs could be readily adapted to a solid electrolyte with capabilities beyond the existing State of Art for specific applications such as In Situ Resource Utilization, lunar fuel cell power systems, or regenerative fuel cell energy storage systems. As Commercial Lunar Payload Services (CLPS) companies have identified primary fuel cell power systems as a required technology, it would be helpful to ensure that there are options available that could survive the lunar night when off-line without active thermal control. This would enable a longer period between missions to re-fuel and recover the electrochemical system.

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Scope Title
Control of Scattered Starlight with Coronagraphs and Starshades

Scope Description

This subtopic addresses the unique problem of imaging and spectroscopic characterization of faint astrophysical objects that are located within the obscuring glare of much brighter stellar sources. Examples include planetary systems beyond our own, the detailed inner structure of galaxies with very bright nuclei, binary star formation, and stellar evolution. Contrast ratios of one million to ten billion over an angular spatial scale of 0.05 - 1.5 arcsec are typical of these objects. Achieving a very low background requires control of both scattered and diffracted light. The failure to control either amplitude or phase fluctuations in the optical train severely reduces the effectiveness of starlight cancellation schemes.

This innovative research focuses on advances in coronagraphic instruments, starlight cancellation instruments, and potential occulting technologies that operate at visible and near infrared wavelengths. The ultimate application of these instruments is to operate in space as part of a future observatory mission concepts such as the Habitable Exoplanet Observatory (HabEx) and Large Ultraviolet Optical Infrared Surveyor (LUVOIR). Measurement techniques include imaging, photometry, spectroscopy, and polarimetry. There is interest in component development and innovative instrument design, as well as in the fabrication of subsystem devices to include, but not limited to, the following areas:

Starlight Suppression Technologies:

• Hybrid metal/dielectric and polarization apodization masks for diffraction control of phase and amplitude for coronagraph scaled starshade experiments.
• Low-scatter, low-reflectivity, sharp, flexible edges for control of solar scatter in starshades.
• Low-reflectivity coatings for flexible starshade optical shields.
• Systems to measure spatial optical density, phase inhomogeneity, scattering, spectral dispersion, thermal variations, and to otherwise estimate the accuracy of high-dynamic range apodizing masks.
• Methods to distinguish the coherent and incoherent scatter in a broadband speckle field.

Wavefront Measurement and Control Technologies:

• Small stroke, high precision, deformable mirrors and associated driving electronics scalable to 10,000 or more actuators (both to further the state-of-the-art towards flight-like hardware and to explore novel concepts). Multiple deformable mirror technologies in various phases of development and processes are encouraged to ultimately improve the state-of-the-art in deformable mirror technology. Process improvements are needed to improve repeatability, yield, and performance precision of current devices.
• Multiplexers with ultra-low power dissipation for electrical connection to deformable mirrors
• Low-order wavefront sensors for measuring wavefront instabilities to enable real-time control and post-processing of aberrations.
• Thermally and mechanically insensitive optical benches and systems.

Optical Coating and Measurement Technologies:

• Instruments capable of measuring polarization cross-talk and birefringence to parts per million.
• Polarization-insensitive coatings for large optics.
• Methods to measure the spectral reflectivity and polarization uniformity across large optics.
• Methods to apply carbon nanotube coatings on the surfaces of the coronagraphs for broadband suppression from visible to near infrared (NIR).

References

See SPIE conference papers and articles published in the Journal of Astronomical Telescopes and Instrumentation on high contrast coronagraphy, segmented coronagraph design and analysis, and starshades.

Websites:

• Exoplanet Exploration - Planets Beyond Our Solar System: https://exoplanets.jpl.nasa.gov
• Exoplanet Exploration Program: https://exoplanets.nasa.gov/exep/
• Goddard Space Flight Center: https://www.nasa.gov/goddard

Expected TRL or TRL range at completion of the project: 3 to 5

Desired Deliverables of Phase II

Prototype, Analysis, Hardware, Research

Desired Deliverables Description

This subtopic solicits proposals to develop components that improve the footprint, robustness, power consumption, reliability, and wavefront quality of high-contrast, low-temporal bandwidth, adaptive optics systems. These include ASIC drivers that easily integrate with the deformable mirrors, improved connectivity technologies, as well as high-actuator count deformable mirrors  with high-quality, ultrastable wavefronts.

It also seeks coronagraph masks that can be tested in ground-based high-contrast testbeds in place at a number of institutions, as well as devices to measure the masks to inform optical models. The masks include transmissive scalar, polarization-dependent, and spatial apodizing masks including those with extremely low reflectivity regions that allow them to be used in reflection.

The subtopic seeks samples of optical coatings that reduce polarization and can be applied to large optics, and methods and instruments to characterize them over large optical surfaces.

Finally, for starshades, the subtopic seeks low reflectivity and potentially diffraction-controlling edges that minimize scattered sunlight while also remaining robust to handling and cleaning.  Low-reflectivity optical coatings that can be applied to the surfaces for the large (hundreds of square meters) optical shield are also desired.

State of the Art and Critical Gaps

Coronagraphs have been demonstrated to achieve high contrast in moderate bandwidth in laboratory environments. Starshades will enable even deeper contrast over broader bands but to date have demonstrated deep contrast in narrow band light. The extent to which the telescope optics will limit coronagraph performance is a function of the quality of the optical coating and the ability to control polarization over the full wavefront. Neither of these technologies is well characterized at levels required for 1e10 contrast. Wavefront control using deformable mirrors is critical. Controllability and stability to picometer levels is required. To date, deformable mirrors have been up to the task of providing contrast approaching 1e10, but they require thousands of wires, and overall wavefront quality and stroke remain concerns.

Relevance / Science Traceability

These technologies are directly applicable to the Wide Field Infrared Survey Telescope (WFIRST), coronagraph instrument (CGI), and the HabEx and LUVOIR concept studies.

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Scope Title

Optical Components and Systems for Large Telescope Missions

Scope Description

To accomplish NASA’s high-priority science at all levels (flagship, probe, Medium-Class Explorers (MIDEX), Small Explorers (SMEX), rocket and balloon) requires low-cost, ultra-stable, normal incidence mirror systems with low mass-to-collecting area ratios. Where a mirror system is defined as the mirror substrate, supporting structure, and associated actuation and thermal management systems. After performance, the most important metric for an advanced optical system is affordability or areal cost (cost per square meter of collecting aperture). Current normal incidence space mirrors cost $4 million to $6 million per square meter of optical surface area. This research effort seeks to improve the performance of advanced precision optical components while reducing their cost by 5 to 50 times, to between $100K/m2 to $1M/m2.

Specific metrics are defined for each wavelength application region:

Aperture Diameter for all wavelengths, except Far-IR

• Monolithic: 1 to 8 meters
• Segmented: 3 to 20 meters

For UV/Optical

• Areal Cost < $500K/m2
• Wavefront Figure < 5 nm RMS (via passive design or active deformation control)
• Wavefront Stability < 10 pm/10 min
• First Mode Frequency 60 to 500 Hz
• Actuator Resolution < 1 nm RMS
• Optical Path-length Stability < 1 pm/10,000 seconds for precision metrology
• Areal density < 15 kg/m2 (< 35 kg/m2 with backplane)
• Operating Temperature Range of 250 to 300K

For Far-IR

• Aperture diameter 1 to 4 m (monolithic), or 5 to 10 m (segmented)
• Telescope diffraction-limited at <30 microns at operating temperature 4 K
• Cryo-Deformation < 100 nm RMS
• Areal cost < $500K/m2
• Production rate > 2 m2 per month
• Areal density < 15 kg/m2 (< 40 kg/m2 with backplane)
• Thermal conductivity at 4 K > 2 W/m*K
• Survivability at temperatures ranging from 315 K to 4 K

For EUV

• Surface Slope < 0.1 micro-radian

Also needed is ability to fully characterize surface errors and predict optical performance.

Proposals must show an understanding of one or more relevant science needs, and present a feasible plan to develop the proposed technology for infusion into a NASA program: sub-orbital rocket or balloon; competed SMEX or MIDEX; or, Decadal class mission. Successful proposals will demonstrate an ability to manufacture, test and control ultra-low-cost optical systems that can meet science performance requirements and mission requirements (including processing and infrastructure issues). Material behavior, process control, active and/or passive optical performance, and mounting/deploying issues should be resolved and demonstrated.

References

The Habitable Exoplanet Imager (HabEx) and Large UVOIR (LUVOIR) space telescope studies are developing concepts for UVOIR space telescopes for exoEarth discovery and characterization, exoplanet science, general astrophysics and solar system astronomy. The HabEx Interim Report is available at: https://www.jpl.nasa.gov/habex/documents/. \ The LUVOIR Interim Report is available at: https://asd.gsfc.nasa.gov/luvoir/.

The Origins Space Telescope (OST) is a single-aperture telescope concept for the Far-Infrared Surveyor mission described in the NASA Astrophysics Roadmap, "Enduring Quests, Daring Visions: NASA Astrophysics in the Next Three Decades": https://smd-prod.s3.amazonaws.com/science-pink/s3fs-public/atoms/files/secure-Astrophysics_Roadmap_2013_0.pdf(link is external).

The OST mission is described on the website: https://origins.ipac.caltech.edu(link is external).

The Space Infrared Interferometric Telescope (SPIRIT) and its optical system requirements are described on the website: https://asd.gsfc.nasa.gov/cosmology/spirit/.

LISA (Laser Interferometer Space Antenna) mission description: https://lisa.nasa.gov/.

Expected TRL or TRL range at completion of the project: 3 to 5

Desired Deliverables of Phase II

Prototype, Hardware, Research

Desired Deliverables Description

An ideal Phase 1 deliverable would be a precision optical system of at least 0.25 meters; or a relevant sub-component of a system; or a prototype demonstration of a fabrication, test or control technology leading to a successful Phase 2 delivery; or a reviewed preliminary design and manufacturing plan which demonstrates feasibility. While detailed analysis will be conducted in Phase 2, the preliminary design should address how optical, mechanical (static and dynamic) and thermal designs and performance analysis will be done to show compliance with all requirements. Past experience or technology demonstrations which support the design and manufacturing plans will be given appropriate weight in the evaluation.

An ideal Phase 2 project would further advance the technology to produce a flight-qualifiable optical system greater than 0.5 meters or relevant sub-component (with a TRL in the 4 to 5 range); or a working fabrication, test or control system. Phase 1 and Phase 2 mirror system or component deliverables would be accompanied by all necessary documentation, including the optical performance assessment and all data on processing and properties of its substrate materials. A successful mission oriented Phase 2 would have a credible plan to deliver for the allocated budget a fully assembled and tested telescope assembly which can be integrated into the potential mission; and, demonstrate an understanding of how the engineering specifications of their system meets the performance requirements and operational constraints of the mission (including mechanical and thermal stability analysis).

State of the Art and Critical Gaps

Current normal incidence space mirrors cost $4 million to $6 million per square meter of optical surface area. This research effort seeks a cost reduction for precision optical components by 5 to 50 times, to between $100K/m2 to $1M/m2.

Relevance / Science Traceability

S2.03 primary supports potential Astrophysics Division missions. S2.03 has made optical systems in the past for potential balloon experiments. Future potential Decadal missions include LISA, Habitable Exoplanet Observatory (HabEx), Large UV/Optical/Near-IR Surveyor (LUVOIR) and the Origins Space Telescope (OST).

 

Scope Title
Balloon Planetary Telescope

Scope Description

Astronomy from a stratospheric balloon platform offers numerous advantages for planetary science. At typical balloon cruise altitudes (100,000 to 130,000 ft.), 99%+ of the atmospheric is below the balloon and the attenuation due to the remaining atmosphere is small, especially in the near ultraviolet band and in the infrared bands near 2.7 and 4.25 µm. The lack of atmosphere nearly eliminates scintillation and allows the resolution potential of relatively large optics to be realized, and the small amount of atmosphere reduces scattered light and allows observations of brighter objects even during daylight hours.

For additional discussion of the advantages of observations from stratosphere platforms, refer to “Planetary Balloon-Based Science Platform Evaluation and Program Implementation - Final Report,” Dankanich et.al. (Available from https://ntrs.nasa.gov/, search for "NASA/TM-2016-218870")

To perform Planetary Science requires a 1-meter class telescope 500 nm diffraction limited performance or Primary Mirror System that can maintain < 10 nm rms surface figure error for elevation angles ranging from 0 to 60 degrees over a temperature range from 220K to 280K.

Phase I will produce a preliminary design and report including initial design requirements such as wave-front error budget, mass allocation budget, structural stiffness requirements, etc., trade studies performed and analysis that compares the design to the expected performance over the specified operating range. Development challenges shall be identified during phase I including trade studies and challenges to be addressed during Phase II with subsystem proof of concept demonstration hardware. If Phase II can only produce a sub-scale component, then it should also produce a detailed final design, including final requirements (wave-front error budget, mass allocation, etc) and performance assessment over the specified operating range.

Additional information about Scientific Balloons can be found at https://www.csbf.nasa.gov/docs.html.

Telescope Specifications:

• Diameter > 1 meter
• System Focal Length 14 meter (nominal)
• Diffraction Limit < 500 nm
• Mass < 300 kg
• Shock 10G without damage
• Elevation 0 to 60 degrees
• Temperature 220 to 280 K

Primary Mirror Assembly Specifications:

• Diameter > 1 meter
• Radius of Curvature 3 meters (nominal)
• Surface Figure Error < 10 nm rms
• Mass < 150 kg
• Shock 10G without damage
• Elevation 0 to 60 degrees
• Temperature 220 to 280 K

References

For additional discussion of the advantages of observations from stratosphere platforms, refer to “Planetary Balloon-Based Science Platform Evaluation and Program Implementation - Final Report,” Dankanich et.al. (Available from https://ntrs.nasa.gov/, search for "NASA/TM-2016-218870")

Expected TRL or TRL range at completion of the project: 3 to 5

Desired Deliverables of Phase II

Prototype, Analysis, Hardware

Desired Deliverables Description

If Phase II can only produce a sub-scale component, then it should also produce a detailed final design, including final requirements (wave-front error budget, mass allocation, etc.) and performance assessment over the specified operating range.

State of the Art and Critical Gaps

To perform Planetary Science requires a 1-meter class telescope 500 nm diffraction limited performance or Primary Mirror System that can maintain < 10 nm rms surface figure error for elevation angles ranging from 0 to 60 degrees over a temperature range from 220K to 280K.
Significant science returns may be realized through observations in the 300 nm to 5 μm range.
Current SOA (State of the Art) mirrors made from Zerodur or ULE for example require light weighting to meet balloon mass limitations, and cannot meet diffraction limited performance over the wide temperature range due to the coefficient of thermal expansion limitations.

Relevance / Science Traceability

From “Vision and Voyages for Planetary Science in the Decade 2013-2022”:

• Page 22, Last Paragraph of NASA Telescope Facilities within the Summary Section:
  Balloon- and rocket-borne telescopes offer a cost-effective means of studying planetary bodies at wavelengths inaccessible from the ground.6 Because of their modest costs and development times, they also provide training opportunities for would-be developers of future spacecraft instruments. Although NASA’s Science Mission Directorate regularly flies balloon missions into the stratosphere, there are few funding opportunities to take advantage of this resource for planetary science, because typical planetary grants are too small to support these missions. A funding line to promote further use of these suborbital observing platforms for planetary observations would complement and reduce the load on the already oversubscribed planetary astronomy program.
• Page 203, 5th paragraph, Section titled Earth and Space-Based Telescopes:
  Significant planetary work can be done from balloon-based missions flying higher than 45,000 ft. This altitude provides access to electromagnetic radiation that would otherwise be absorbed by Earth’s atmosphere and permits high-spatial-resolution imaging unaffected by atmospheric turbulence. These facilities offer a combination of cost, flexibility, risk tolerance, and support for innovative solutions that is ideal for the pursuit of certain scientific opportunities, the development of new instrumentation, and infrastructure support. Given the rarity of giant-planet missions, these types of observing platforms (high-altitude telescopes on balloons and sounding rockets) can be used to fill an important data gap.154, 155,156.

Potential Advocates include Planetary Scientists at GSFC, APL, and Southwest Research Institute, etc. The NASA Balloon Workshop.

Potential Projects Gondola for High Altitude Planetary Science (GHAPS).

 

Scope Title
Large UV/Optical (LUVOIR) and Habitable Exoplanet (HabEx) Missions

Scope Description

Potential UV/Optical missions require 4 to 16 meter monolithic or segmented primary mirrors with < 5 nm RMS surface figures. Active or passive alignment and control is required to achieve system level diffraction limited performance at wavelengths less than 500 nm (< 40 nm RMS wavefront error, WFE). Additionally, potential Exoplanet mission, using an internal coronagraph, requires total telescope wavefront stability on order of 10 pico-meters RMS per 10 minutes. This stability specification places severe constraints on the dynamic mechanical and thermal performance of 4 meter and larger telescope. Potential enabling technologies include: active thermal control systems, ultra-stable mirror support structures, athermal telescope structures, athermal mirror struts, ultra-stable low CTE/high-stability joints, and vibration compensation.

Mirror areal density depends upon available launch vehicle capacities to Sun-Earth L2 (i.e. 15 kg/m2 for a 5 m fairing EELV vs. 150 kg/m2 for a 10 m fairing SLS). Regarding areal cost, a good goal is to keep the total cost of the primary mirror at or below $100M. Thus, an 8-m class mirror (with 50 m2 of collecting area) should have an areal cost of less than $2M/m2. And, a 16-m class mirror (with 200 m2 of collecting area) should have an areal cost of less than $0.5M/m2.

Key technologies to enable such a mirror include new and improved:

• Mirror substrate materials and/or architectural designs
• Processes to rapidly fabricate and test UVO quality mirrors
• Mirror support structures, joints and mechanisms that are athermal or zero CTE at the desired scale
• Mirror support structures, joints and mechanisms that are ultra-stable at the desired scale
• Mirror support structures with low-mass that can survive launch at the desired scale
• Mechanisms and sensors to align segmented mirrors to < 1 nm RMS precisions
• Thermal control (< 1 mK) to reduce wavefront stability to < 10 pm RMS per 10 min
• Dynamic isolation (> 140 dB) to reduce wavefront stability to < 10 pm RMS per 10 min

Also needed is ability to fully characterize surface errors and predict optical performance via integrated opto-mechanical modeling.

Potential solutions for substrate material/architecture include, but are not limited to: ultra-uniform low CTE glasses, silicon carbide, nanolaminates or carbon-fiber reinforced polymer. Potential solutions for mirror support structure material/architecture include, but are not limited to: additive manufacturing, nature inspired architectures, nanoparticle composites, carbon fiber, graphite composite, ceramic or SiC materials, etc. Potential solutions for new fabrication processes include, but are not limited to: additive manufacture, direct precision machining, rapid optical fabrication, roller embossing at optical tolerances, slumping or replication technologies to manufacture 1 to 2 meter (or larger) precision quality components. Potential solutions for achieving the 10 pico-meter wavefront stability include, but are not limited to: metrology, passive, and active control for optical alignment and mirror phasing; active vibration isolation; metrology, passive, and active thermal control.

References

The Habitable Exoplanet Imager (HabEx) and Large UVOIR (LUVOIR) space telescope studies are developing concepts for UVOIR space telescopes for exoEarth discovery and characterization, exoplanet science, general astrophysics and solar system astronomy. The HabEx Interim Report is available at: https://www.jpl.nasa.gov/habex/pdf/interim_report.pdf. The LUVOIR Interim Report is available at: https://asd.gsfc.nasa.gov/luvoir/.

The Origins Space Telescope (OST) is a single-aperture telescope concept for the Far-Infrared Surveyor mission described in the NASA Astrophysics Roadmap, "Enduring Quests, Daring Visions: NASA Astrophysics in the Next Three Decades" (https://smd-prod.s3.amazonaws.com/science-pink/s3fs-public/atoms/files/secure-Astrophysics_Roadmap_2013_0.pdf(link is external)).

The OST mission is described on the website https://origins.ipac.caltech.edu(link is external).

The Space Infrared Interferometric Telescope (SPIRIT) and its optical system requirements are described on the website https://asd.gsfc.nasa.gov/cosmology/spirit/.

Expected TRL or TRL range at completion of the project: 2 to 4

Desired Deliverables of Phase II

Analysis, Hardware, Software, Research

Desired Deliverables Description

An ideal Phase 1 deliverable would be a precision optical system of at least 0.25 meters; or a relevant sub-component of a system; or a prototype demonstration of a fabrication, test or control technology leading to a successful Phase 2 delivery; or a reviewed preliminary design and manufacturing plan which demonstrates feasibility. While detailed analysis will be conducted in Phase 2, the preliminary design should address how optical, mechanical (static and dynamic) and thermal designs and performance analysis will be done to show compliance with all requirements. Past experience or technology demonstrations which support the design and manufacturing plans will be given appropriate weight in the evaluation.

An ideal Phase 2 project would further advance the technology to produce a flight-qualifiable optical system greater than 0.5 meters or relevant sub-component (with a TRL in the 4 to 5 range); or a working fabrication, test or control system. Phase 1 and Phase 2 mirror system or component deliverables would be accompanied by all necessary documentation, including the optical performance assessment and all data on processing and properties of its substrate materials. A successful mission oriented Phase 2 would have a credible plan to deliver for the allocated budget a fully assembled and tested telescope assembly which can be integrated into the potential mission; and, demonstrate an understanding of how the engineering specifications of their system meets the performance requirements and operational constraints of the mission (including mechanical and thermal stability analysis).

State of the Art and Critical Gaps

Hubble at 2.4m is the SOA.

Relevance / Science Traceability

S2.03 primary supports potential Astrophysics Division missions. S2.03 has made optical systems in the past for potential balloon experiments. Future potential Decadal missions include LISA, Habitable Exoplanet Observatory (HabEx), Large UV/Optical/Near-IR Surveyor (LUVOIR) and the Origins Space Telescope (OST).

 

Scope Title
NIR LIDAR Beam Expander Telescope

Scope Description

Potential airborne coherent LIDAR missions need compact 15-cm diameter 20X magnification beam expander telescopes. Potential space based coherent LIDAR missions need at least 50-cm 65X magnification beam expander telescopes. Candidate coherent LIDAR systems (operating with a pulsed 2-micrometer laser) have a narrow, almost diffraction limited field of view, close to 0.8 lambda/D half angle. Aberrations, especially spherical aberration, in the optical telescope can decrease the signal. Additionally, the telescope beam expander should maintain the laser beam’s circular polarization. The incumbent telescope technology is a Dahl-Kirkham beam expander. Technology advance is needed to make the beam expander more compact with less mass while retaining optical performance, and to demonstrate the larger diameter. Additionally, technology for non-moving scanning of the beam expander output is needed.

References

NRC Decadal Surveys at: http://sites.nationalacademies.org/DEPS/ESAS2017/index.htm(link is external).

https://smd-prod.s3.amazonaws.com/science-pink/s3fs-public/atoms/files/Weather_Focus_Area_Workshop_Report_2015_0.pdf(link is external).

A. K. DuVivier, J. J. Cassano, S. Greco and G. D. Emmitt, 2017, “A Case Study of Observed and Modeled Barrier Flow in the Denmark Strait in May 2015” Monthly Weather Review 145, 2385 – 2404 (2017): (https://doi.org/10.1175/MWR-D-16-0386.1(link is external)). See also Supplemental Material: http://dx.doi.org/10.1175/MWR-D-16-0386.s1(link is external)

M. J. Kavaya, J. Y. Beyon, G. J. Koch, M. Petros, P. J. Petzar, U. N. Singh, B. C. Trieu, and J. Yu, “The Doppler Aerosol Wind Lidar (DAWN) Airborne, Wind-Profiling, Coherent-Detection Lidar System: Overview and Preliminary Flight Results,” J. of Atmospheric and Oceanic Technology 34 (4), 826-842 (2014), https://doi.org/10.1175/JTECH-D-12-00274.1(link is external)

Scott A. Braun, Ramesh Kakar, Edward Zipser, Gerald Heymsfield, Cerese Albers, Shannon Brown, Stephen L. Durden, Stephen Guimond, Jeffery Halverson, Andrew Heymsfield, Syed Ismail, Bjorn Lambrigtsen, Timothy Miller, Simone Tanelli, Janel Thomas , and Jon Zawislak, “NASA’s Genesis and Rapid Intensification Processes (GRIP) Field Experiment,” Bull. Amer. Meteor. Soc. (BAMS) 94(3), 345-363 (2013), https://doi.org/10.1175/BAMS-D-11-00232.1(link is external)

Expected TRL or TRL range at completion of the project: 3 to 4

Desired Deliverables of Phase II

Prototype, Analysis, Hardware, Research

Desired Deliverables Description

A detailed design or a small prototype or a full-sized beam expander.

State of the Art and Critical Gaps

The current SOA is a COTS beam expander with a 15-cm diameter primary mirror, a heavy aluminum structure, an Invar rod providing thermally insensitive primary-to-secondary mirror separation, and a manually adjustable and lockable variable focus setting by changing the mirror separation. Critical gaps include 1) a 50-70 cm diameter primary mirror beam expander that features near-diffraction limited performance, low mass design, minimal aberrations with an emphasis on spherical, characterization of the polarization changes vs. beam cross section assuming input circular polarization, a lockable electronic focus adjustment, both built-in and removable fiducial aids for aligning the input laser beam to the optical axis, and a path to space qualification; and 2) a 15-cm diameter primary mirror beam expander with the same features for airborne coherent lidar systems.

Relevance / Science Traceability

Science Mission Directorate (SMD) desires both an airborne coherent-detection wind-profiling lidar systems and a space-based wind measurement. The space mission has been recommended to SMD by both the 2007 and 2017 earth science Decadal Surveys. SMD has incorporated the wind lidar mission in its planning and has named it "3-D Winds". SMD recently held the Earth Venture Suborbital competition for 5-years of airborne science campaigns. The existing coherent wind lidar at Langley, DAWN, was included in three proposals which are under review. Furthermore, SMD is baselining DAWN for a second CPEX-type airborne science campaign, and for providing cal/val assistance to the ESA AEOLUS space mission. DAWN flies on the DC-8 and it is highly desired to fit DAWN on other NASA and NOAA aircraft. DAWN needs to lower its mass for several of the aircraft, and a low-mass telescope retaining the required performance is needed. Additionally, an electronic remote control of telescope focus is needed to adapt to aircraft cruise altitude and weather conditions during science flights.

 

Scope Title
Fabrication, Test and Control of Advanced Optical Systems

Scope Description

Future UV/Optical/NIR telescopes require mirror systems that are very precise and ultra-stable.

Regarding precision, this subtopic encourages proposals to develop technology which makes a significant advance the ability to fabricate and test an optical system.

One area of current emphasis is the ability to non-destructively characterize CTE homogeneity in 4-m class Zerodur and 2-m class ULE mirror substrates to an uncertainty of 1 ppb/K and a spatial sampling of 100 x 100.  This characterization capability is needed to select mirror substrates before they undergo the expense of turning them into a light-weight space mirror.

Regarding stability, to achieve high-contrast imaging for exoplanet science using a coronagraph instrument, systems must maintain wavefront stability to < 10 pm RMS over intervals of ~10 minutes during critical observations. The ~10-minute time period of this stability is driven by current wavefront sensing and control techniques that rely on stellar photons from the target object to generate estimates of the system wavefront. This subtopic aims to develop new technologies and techniques for wavefront sensing, metrology, and verification and validation of optical system wavefront stability.

Current methods of wavefront sensing include image-based techniques such as phase retrieval, focal-plane contrast techniques such as electric field conjugation and speckle nulling, and low-order and out-of-band wavefront sensing that use non-science light rejected by the coronagraph to estimate drifts in the system wavefront during observations. These techniques are limited by the low stellar photon rates of the dim objects being observed (~5 - 11 Vmag), leading to 10s of minutes between wavefront control updates.

New methods may include: new techniques of using out-of-band light to improve sensing speed and spatial frequency content, new control laws incorporating feedback and feedforward for more optimal control, new algorithms for estimating absolute and relative wavefront changes, and the use of artificial guide stars for improved sensing signal to noise ratio and speed.

Current methods of metrology include edge sensors (capacitive, inductive, or optical) for maintaining segment cophasing, and laser distance interferometers for absolute measurement of system rigid body alignment. Development of these techniques to improve sensitivity, speed, and component reliability is desired. Low power, high-reliability electronics are also needed.

Finally, metrology techniques for system verification and validation at the picometer level during integration and test (I&T) are needed. High speed spatial and speckle interferometers are currently capable of measuring single-digit picometer displacements and deformations on small components in controlled environments. Extension of these techniques to large-scale optics and structures in typical I&T environments is needed.

References

The HabEx Interim Report is available at: https://www.jpl.nasa.gov/habex/pdf/interim_report.pdf. The LUVOIR Interim Report is available at: https://asd.gsfc.nasa.gov/luvoir/resources/docs/LUVOIR_Interim_Report_Fi....

Expected TRL or TRL range at completion of the project: 2 to 4

Desired Deliverables of Phase II

Analysis, Hardware, Software, Research

Desired Deliverables Description

An ideal Phase 1 deliverable would be a prototype demonstration of a fabrication, test or control technology leading to a successful Phase 2 delivery; or a reviewed preliminary design and manufacturing plan which demonstrates feasibility. While detailed analysis will be conducted in Phase 2, the preliminary design should address how optical, mechanical (static and dynamic) and thermal designs and performance analysis will be done to show compliance with all requirements. Past experience or technology demonstrations which support the design and manufacturing plans will be given appropriate weight in the evaluation.

An ideal Phase 2 project would further advance the technology to produce a flight-qualifiable relevant sub-component (with a TRL in the 4 to 5 range); or a working fabrication, test or control system. Phase 1 and Phase 2 mirror system or component deliverables would be accompanied by all necessary documentation, including the optical performance assessment and all data on processing and properties of its substrate materials. A successful mission oriented Phase 2 would have a credible plan to deliver for the allocated budget a fully assembled and tested telescope assembly which can be integrated into the potential mission; and, demonstrate an understanding of how the engineering specifications of their system meets the performance requirements and operational constraints of the mission (including mechanical and thermal stability analysis).

State of the Art and Critical Gaps

Wavefront sensing using star images, including dispersed-fringe and phase retrieval methods, is at TRL 6, qualified for space by JWST. Wavefront sensing and control for coronagraphs, including electric field conjugation and Low-Order WF Sensing (LOWFS) is at TRL4, and is being developed and demonstrated by WFIRST/CGI.

Laser distance interferometers for point-to-point measurements with accuracies from nanometers to picometers have been demonstrated on the ground by the Space Interferometry Mission and other projects, and on orbit by the Lisa Pathfinder and Grace Follow-On mission. Application to telescope alignment metrology has been demonstrated on testbeds, to TRL4 for nanometer accuracy. Picometer accuracy for telescopes awaits demonstration.

Edge sensors are in use on segmented ground telescopes, but not yet on space telescopes. New designs are needed to provide picometer sensitivity and millimeter range in a space qualified package.

Higher-order WFS for coronagraphs using out-of-band light is beginning development, with data limited to computer simulations. Such techniques are best used

Relevance / Science Traceability

These technologies are enabling for coronagraph-equipped space telescopes, segmented space telescopes, and others that utilize actively controlled optics. The LUVOIR and HabEx mission concepts currently under study provide good examples.

 

Scope Title
Optical Components and Systems for potential Infrared/Far-IR missions

Scope Description

The Far-IR Surveyor Mission described in NASA's Astrophysics Roadmap, "Enduring Quests, Daring Visions":

In the context of subtopic S2.03, the challenge is to take advantage of relaxed tolerances stemming from a requirement for long wavelength (30 micron) diffraction-limited performance in the fully-integrated optical telescope assembly to minimize the total mission cost through innovative design and material choices and novel approaches to fabrication, integration, and performance verification.

The Far-IR Surveyor is a cryogenic far-infrared mission, which could be either a large single-aperture telescope or an interferometer. There are many common and a few divergent optical system requirements between the two architectures.

Common requirements:

• Telescope operating temperature ~4 K
• Telescope diffraction-limited at 30 microns at the operating temperature
• Mirror survivability at temperatures ranging from 315 K to 4 K
• Mirror substrate thermal conductivity at 4 K > 2 W/m*K
• Zero or low CTE mismatch between mirror substrate and backplane

Divergent requirements:

• Large single-aperture telescope:
  • Segmented primary mirror, circular or hexagonal
  • Primary mirror diameter 5 to 10 m
  • Possible 3 dof (tip, tilt and piston) control of mirror segments on orbit
  • Interferometer:
    • Monolithic primary mirrors
    • Afocal, off-axis telescope design
    • Primary mirror diameter 1 to 4 m

Success metrics:

• Areal cost < $500K/m2
• Areal density < 15 kg/m2 (< 40 kg/m2 with backplane)
• Production rate > 2 m2 per month
• Short time span for optical system integration and test

References

The Far-Infrared Surveyor is described in NASA's Astrophysics Roadmap, "Enduring Quests, Daring Visions," which can be downloaded from https://smd-prod.s3.amazonaws.com/science-pink/s3fs-public/atoms/files/secure-Astrophysics_Roadmap_2013_0.pdf(link is external).

Program Annual Technology Reports (PATR) can be downloaded from the NASA PCOS/COR Technology Development website at https://apd440.gsfc.nasa.gov/technology/.

Expected TRL or TRL range at completion of the project: 3 to 5

Desired Deliverables of Phase II

Prototype, Hardware, Research

Desired Deliverables Description

Mirrors or optical systems that demonstrably advance TRL to address the overall challenge described under Scope Description while meeting requirements for a single-aperture or interferometric version of the notional Far-IR Surveyor mission.

State of the Art and Critical Gaps

Current SOA is represented by the Herschel Space Observatory (3.5 m monolith; SiC) and James Webb Space Telescope (6.5 m segmented primary mirror; beryllium).

Relevance / Science Traceability

The technology is relevant to the Far-IR Surveyor mission described in NASA's Astrophysics Roadmap and prioritized in NASA's Program Annual Technology Reports for Cosmic Origins and Physics of the Cosmos. A future NASA far-infrared astrophysics mission will answer compelling questions, such as: How common are life-bearing planets?; How do the conditions for habitability develop during the process of planet formation?; and How did the universe evolve in response to its changing ingredients (build-up of heavy elements and dust over time)? To answer these questions, NASA will need telescopes and interferometers that reach fundamental sensitivity limits imposed by astrophysical background photon noise. Only telescopes cooled to a cryogenic temperature can provide such sensitivity.

Novel approaches to fabrication and test developed for a far-infrared astrophysics mission may be applicable to far-infrared optical systems employed in other divisions of the NASA Science Mission Directorate, or to optical systems designed to operate at wavelengths shorter than the far-infrared.

 

Scope Title
Low-Cost Compact Reflective Telescope for NIR/SWIR Optical Communication

Scope Description

The need exists for a low cost methodology to produce compact (for ex., cubesat-class), scalable, diffraction limited, athermalized, off-axis reflective-type, optics for NIR/SWIR-band communication applications. Typically, specialty optical aperture systems are designed and built as “one-offs” which are inherently high in cost and often out of scope for smaller projects. A Phase I would investigate current compact off-axis reflective designs and develop a trade space to identify the most effective path forward. The work would include a strategy for aperture diameter scalability, athermalization, and low cost fabrication. Detailed optical designs would be developed along with detailed structural, thermal, optical performances (STOP) analyses confirming diffraction limited operation across a wide range of operational disturbances, both structural dynamic and thermal. Commercial of the shelf (COTS) NIR/SWIR optical communication support hardware should be assumed towards an integrated approach, including fiber optics, fast steering mirrors, and applicable detectors. Phase II may follow up with development of prototypes, built at multiple aperture diameters and fidelities.

References

An example of an on-axis design has been utilized in LLCD: https://www.spiedigitallibrary.org/conference-proceedings-of-spie/10563/105630X/NASAs-current-activities-in-free-space-optical-communications/10.1117/12.2304175.full?SSO=1(link is external)

An example of an off-axis design is being developed by JPL for deep space optical comm (DSOC): https://www.spiedigitallibrary.org/conference-proceedings-of-spie/10096/100960V/Discovery-deep-space-optical-communications-DSOC-transceiver/10.1117/12.2256001.full(link is external)

Expected TRL or TRL range at completion of the project: 2 to 4

Desired Deliverables of Phase II

Prototype, Analysis, Hardware

Desired Deliverables Description

Prototype unobscured telescope with the required scale size

State of the Art and Critical Gaps

Currently, the state of the art for reflective optical system for communications applications are:

1)     On-axis or axisymmetric designs are typically used for (space) optical comm and imaging, which inherently are problematic due to the central obscuration.

2)     Off-axis designs provide superior optical performance due to the clear aperture, however, are rarely considered due to complex design, manufacturing, and metrology procedures needed.

Relevance / Science Traceability

Optical Communication enable high data-rate downlink of science data. The initial motivation for this scalable off-axis optical design approach is for bringing high-performance reflective optics within reach of laser communication projects with limited resources. However, this exact optical hardware is applicable for any diffraction limited, athermalized science imaging applications. Any science mission could potentially be able to select from a “catalog” of optical aperture systems that would already have (flight) heritage and reduced risks.

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Scope Title
Components, assemblies, and subsystems for Extreme Precision Radial Velocity Measurements and Detection of Extrasolar Planets

Scope Description

Astronomical spectrographs have proven to be powerful tools for exoplanet searches. When a star experiences periodic motion due to the gravitational pull of an orbiting planet, its spectrum is Doppler-modulated in time. This is the basis for the Precision Radial Velocity (PRV) method, one of the first and most efficient techniques for detecting and characterizing exoplanets. Since spectrographs have their own drifts which must be separated from the periodic Doppler shift, a stable reference is always needed for calibration. Optical Frequency Combs (OFCs) and line-referenced etalons are capable of providing the instrument precision needed for detecting and characterizing Earth-like planets in the Habitable Zone of their Sun-like host stars. While “stellar jitter” (a star’s photospheric velocity contribution to the RV signal) is unavoidable, the contribution to the error budget from Earth’s atmosphere would be eliminated in future space missions. Thus, there is a need to develop robust spectral references with Size, Weight and Power (SWaP) suitable for space qualified operation to calibrate the next generation of high-resolution spectrographs with precision corresponding to < ~1 cm/s over multiple years of observations.

This subtopic solicits proposals to develop cost effective component and subsystem technology for low SWaP, long-lived, robust implementation of radial velocity measurement instruments both on the ground and in space. Research areas of interest include but are not limited to:

• Integrated photonic spectrographs
• PRV spectrograph calibration sources
• High efficiency photonic lanterns
• Advanced fiber scrambling techniques for modal noise reduction
• Software for advanced statistical techniques to mitigate effects of telluric absorption and stellar jitter on RV precision and accuracy

References

Precision Radial Velocity:

• Fischer et al. (2016) State of the Field: Extreme Precision Radial Velocities http://adsabs.harvard.edu/abs/2016PASP..128f6001(link is external)
• Plavchan et al. (2015) Radial Velocity Prospects Current and Future: A White Paper Report prepared by the Study Analysis Group 8 for the Exoplanet Program Analysis Group (ExoPAG) http://adsabs.harvard.edu/abs/2015arXiv150301770P(link is external)
• Plavchan et al. (2019) EarthFinder Probe Mission Concept Study (Final Report): https://smd-prod.s3.amazonaws.com/science-red/s3fs-public/atoms/files/Ea...(link is external)

Photonic Lanterns:

• Gris-Sanchez et al. (2018) Multicore fibre photonic lanterns for precision radial velocity Science: https://academic.oup.com/mnras/article/475/3/3065/4769655(link is external)
• Jvanovic, N. et al. (2012). Integrated photonic building blocks for next-generation astronomical instrumentation I: the multimode waveguide. Optics Express, 20:17029.

Astrocombs:

• Yi, X., et al. (2016) Demonstration of a near-IR line-referenced electro-optical laser frequency comb for precision radial velocity measurements in astronomy. Nature Communications, 7:10436.
• Halverson, S., et al, (2014) "The habitable-zone planet finder calibration system", Proc. SPIE 9147, Ground-based and Airborne Instrumentation for Astronomy V, 91477Z:  https://doi.org/10.1117/12.2054967(link is external)
• Suh, M.-G., et al. (2019) Searching for exoplanets using a microresonator astrocomb. Nature Photonics, 13(1):25–30.
• Obrzud, E., et al. (2019) A Microphotonic Astrocomb. Nature Photonics, 13 (1):31–35.

Nonlinear Waveguides:

• Chang, L., et al. (2018) Heterogeneously integrated GaAs waveguides on insulator for efficient frequency conversion, Laser Photonics Reviews, 12, 1800149: https://doi.org/10.1002/lpor.201800149(link is external)
• Halir, R., et al. (2012) Ultrabroadband supercontinuum generation in a CMOS-compatible platform, Optics letters, 37, 1685: https://di.org/10.1364/OL.37.001685(link is external)

Spectral Flattening:

• Probst, R.A., et al. (2015) Spectrally Flattened, Broadband Astronomical Frequency Combs  https://di.org/10.1364/CLEO_SI.2015.SW4G.7(link is external)

Expected TRL or TRL range at completion of the project: 3 to 5

Desired Deliverables of Phase II

Hardware/software

Desired Deliverables Description

This subtopic solicits proposals to develop cost effective component and subsystem technology for low SWaP, long-lived, robust implementation of radial velocity measurement instruments both on the ground and in space. Research areas of interest include but are not limited to:

• Integrated photonic spectrographs that meet PRV specifications (e.g. wavelength coverage, resolution, throughput, and polarization). These devices should be able to accept multiple fibers - at least two for the science light and simultaneous calibration light source. Ideally, they should be able to include on-chip cross-dispersion to eliminate bulk optics.
• PRV spectrograph calibration sources, particularly optical frequency combs (a.k.a. “astrocombs”) from the UV through the NIR (~350 nm – ~2400 nm) with ~10-30 GHz mode spacing, potentially self-referenced, or line stabilized for Allan Deviation <1E-11 over 100 seconds to years
  • Spectral flattening to provide uniform power across the spectral band covered by the instrument
  • Spectral broadening to obtain wide spectral coverage, preferably octave-spanning to enable self-referencing
  • Integrated photonic solutions including nonlinear waveguides, microresonators or other comb generators, pump lasers, and f-2f beat-note generation
  • Low phase-noise solutions
  • Tunability of comb lines to scan spectrograph detectors for pixel characterization
• Optical etalons with similar requirements for stability as the frequency combs
• High efficiency photonic lanterns
• Advanced fiber scrambling techniques for modal noise reduction
• Software for advanced statistical techniques to mitigate effects of telluric absorption and stellar jitter on RV precision and accuracy.

Proposals should show an understanding of the science needs, as well as present a feasible plan to fully develop the relevant subsystem technologies and to transition into future NASA program(s).

Phase I will emphasize research aspects for technical feasibility, infusion potential into ground or space operations, clear and achievable benefits (e.g., reduction in SWaP and/or cost, improved RV precision), and show a path towards a Phase II proposal. Phase I Deliverables include feasibility and concept of operations of the research topic, simulations and measurements, validation of the proposed approach to develop a given product (TRL 3-4), and a plan for further development of the specific capabilities or products to be performed in Phase II. Early development and delivery of prototype hardware/software is encouraged.

Phase II will emphasize hardware/software development with delivery of specific hardware or software products for NASA targeting demonstration operations at a ground-based telescope in coordination with the lead NASA center. Phase II deliverables include a working prototype or engineering model of the proposed product/platform or software, along with documentation of development, capabilities, and measurements (showing specific improvement metrics), documents and tools as necessary. Proposed prototypes shall demonstrate a path towards a flight-capable platform. Opportunities and plans should also be identified and summarized for potential commercialization or NASA infusion. 

State of the Art and Critical Gaps

The classical bulk optic spectrographs that are traditionally used for PRV science impose architectural constraints due to their large mass and limited optical flexibility. The spectrograph is the single element that if replaced with a photonic alternative could dramatically alter the course of astronomical instrumentation. Integrated Photonic Spectrographs (IPS) are wafer thin devices that could reduce instrument volume by up to three orders of magnitude. Furthermore, high resolving power spectrographs (R~150,000) with simultaneous UV, visible, and NIR coverage and exquisite long-term stability are required for PRV studies. Spectrometers that are fiber-fed with high illumination stability, excellent wavelength calibration, and precise temperature and pressure control represent the immediate future of precision RV measurements.

As spectrograph stability imposes limits on how precisely the Radial Velocity (RV) can be measured, spectral references play a critical role in characterizing and ensuring this precision. Only Laser Frequency Combs (LFCs) and line-referenced Fabry-Pérot etalons are capable of providing the broad spectral coverage and long term (years) stability needed for extreme PRV detection of exoplanets. While both frequency combs and etalons can deliver high precision spectrograph calibration, the former requires relatively complex and sophisticated hardware in the visible portion of the spectrum. Visible band frequency combs for astronomy (a.k.a. astrocombs) were initially based on mode-locked laser comb technology. However, the intrinsic free spectral range of these instruments, 100s of MHz to 1 GHz, is too fine to be resolved by astronomical spectrographs of R~150,000 or less. Thus, mode filtering of comb lines to create a more spectrally sparse calibration grid is necessary. The filtering step introduces complexity and additional sources of instability to the calibration process, as well as instrument assemblies too large in mass and volume for flight.

Commercial fiber laser astrocombs covering 450 - 1400 nm at 25 GHz line spacing and <3 dB intensity variations over the entire bandwidth are available for ground-based astronomical spectrographs and have been developed for HARPS-S and ESPRESSO RV instruments. However, the cost for these systems is often so prohibitive that recent RV spectrograph projects such as CARMENES and Keck Planet Finder either do not use a frequency comb or include it only as a future upgrade, owing to the cost impact on the project.

Alternatively, frequency combs produced by Electro-Optic Modulation (EOM) of a laser source have been demonstrated at observatories for PRV studies in the near-IR. EOM combs produce modes spaced at a RF modulation frequency, typically 10-30 GHz, and are inherently suitable as ground-based astrocombs. Significantly, EOM combs avoid the line filtering step of commercial mode-locked fiber laser combs. Comb frequency stabilization can be accomplished in a variety of ways, including referencing the laser pump source to a molecular absorption feature or another frequency comb. Where octave spanning EOM combs are available, f-2f self-referencing provides the greatest stability.

EOM combs must be spectrally broadened to provide the octave bandwidth necessary for f-2f stabilization for stability traceable to the Standard International (SI) second. This is accomplished through pulse amplification followed by injection into Highly Non-Linear Fiber (HNLF) or nonlinear optical waveguides, but the broadening process is accompanied by multiplication of the optical phase noise from the EOM comb modulation signal and must be optically filtered. Also, at these challenging microwave pulse repetition rates, the pulse duty-cycle requires pulse amplification to 4-5 Watts of average optical power in order to generate the high enough peak intensity needed for nonlinear broadening. This necessitates use of high power, non-telecom amplifiers that are more prone to lifetime issues, making EOM combs not optimal for flight either. It is important to note that very little comb light is actually required on the spectrograph detectors for calibration. In fact, most of the generated comb light must be deliberately attenuated to avoid detector saturation.

Power consumption of the frequency comb calibration system will be a significant driver of mission cost for space-based PRV systems, and motivates the development of a comb system that operates with less than 20 Watts of spacecraft power. Thus, for flight applications, it is highly desirable to develop frequency comb technology with low power consumption, ~10 GHz mode spacing, compact size, broad (octave spanning) spectral grasp across both the visible and NIR, phase noise insensitivity, stability traceable to the definition of the SI second, and very importantly, long life.

Relevance / Science Traceability

The NASA Strategic Plan (2018) and Space Mission Directorate Science Plan (2014) both call for discovery and characterization of habitable Earth analogs and the search for biosignatures on those worlds. These goals were endorsed and amplified upon in the recent National Academy of Science (NAS) Exoplanet Report which emphasized that a knowledge of the orbits and masses is essential to the complete and correct characterization of potentially habitable worlds. PRV measurements are needed to follow up on the transiting worlds discovered by Kepler, K2, and Transiting Exoplanet Survey Satellite (TESS). The interpretation of the transit spectra which James Webb Space Telescope (JWST) will obtain will depend on knowledge of a planet’s surface gravity which comes from its radius (from the transit data) and its mass (from PRV measurements or in some cases Transit Timing Variations). Without knowledge of a planet's mass, the interpretation of its spectrum is subject to many ambiguities.

These ambiguities will only be exacerbated for the direct imaging missions such as the proposed Habitable Exoplanet Observatory (HabEx) and Large Ultraviolet Optical Infrared Surveyor (LUVOIR) flagships which will obtain spectra of Earth analogs around a few tens to hundreds of stars. Even if a radius can be inferred from the planet's brightness and an estimate of its albedo, the lack of a dynamical mass precludes any knowledge of the planet's density, bulk composition, and surface gravity which are needed to determine, for example, absolute gas column densities. Moreover, a fully characterized orbit is challenging to determine from just a few direct images and may even be confused in the presence of multiple planets. Is a planet in a highly eccentric orbit habitable or not? Only dynamical (PRV) measurements can provide such information. Thus, highly precise and highly stable PRV measurements are absolutely critical to the complete characterization of habitable worlds.

The NAS report also noted that measurements from space might be a final option if the problem of telluric contamination cannot be solved. The Earth’s atmosphere will limit precise radial velocity measurements to ~10 cm/s at wavelengths longer than ~700 nm and greater than 30 cm/s at >900 nm, making it challenging to mitigate the effects of stellar activity without a measurement of the color dependence due to stellar activity in the PRV time series. A space-based PRV mission, such as has been suggested in the NASA EarthFinder mission concept study, may be necessary. If so, the low SWaP technologies developed under this SBIR program could help enable space-based implementations of the PRV method.

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Scope Title
Hot Structure Technology for Aerospace Vehicles

Scope Description

This subtopic encompasses the development of reusable hot structure technology for structural components exposed to extreme heating environments on aerospace vehicles. A hot structure system is a multi-functional structure that can reduce or eliminate the need for a separate thermal protection system (TPS) or active cooling system. The potential advantages of using a hot structure system in place of a TPS with underlying cool structure are: reduced mass, increased mission capability, such as reusability, improved aerodynamics, improved structural efficiency and increased ability to inspect the structure.  Hot structure is an enabling technology for reusability between missions or mission phases, such as aerocapture followed by entry, and has been used in prior NASA programs (Space Shuttle Orbiter, Hyper-X, and X-37) on control surfaces and wing leading edges, as well as in Department of Defense programs.  Additionally, the development of hot structure technology for combustion-device liquid rocket engine propulsion systems is of great interest.

This subtopic seeks to develop innovative low-cost, damage tolerant, reusable and lightweight hot structure technology applicable to aerospace vehicles exposed to extreme temperatures between 1093° to 2204°C (2000° to 4000°F). These aerospace vehicle applications are unique in requiring the hot structure to carry primary structure vehicle loads and to be reusable after exposure to extreme temperatures during atmospheric entry and/or liquid rocket engine firings. The material systems of interest for use in developing hot structure technology include:  advanced carbon-carbon (C-C) materials, ceramic matrix composites (CMC’s), or advanced high-temperature refractory metals.  Potential applications of hot structure technology include: primary load-carrying aeroshell structures, control surfaces, leading edges, and propulsion system components (such as hot gas valves, combustion chambers, and passively- or actively-cooled nozzle extensions).

Proposals should present approaches to address the current need for improvements in operating temperature capability, toughness/durability, reusability and material system properties. Focus areas should address one or more of the following:

• Improvements in manufacturing processes and/or material designs to achieve repeatable and uniform material properties that should be scalable to actual vehicle components: specifically, material property data obtained from flat-panel test coupons should represent the properties of prototype and flight test articles.
• Material/structural architectures and multifunctional systems providing significant improvements over typical 2D inter-laminar mechanical properties while maintaining in-plane and thermal properties when compared to state-of-the-art C-C or CMC materials.  Examples include:  incorporating through-the-thickness stitching, braiding or 3D woven preforms.
• Functionally-graded manufacturing approaches to optimize oxidation protection, damage tolerance and structural efficiency, in an integrated hot structure concept that extends performance for multiple cycles up to 2204°C (4000°F).
• Manufacturing process methods that enable a significant reduction in the time required to fabricate materials and components.  There is a great need to reduce processing time for hot structure materials and components -- current state-of-the-art fabrication times are often in the range of 6 to 12 months, which can limit the use of such materials.  Approaches enabling reduced manufacturing times should not, however, lead to significant reductions in material properties.

Under this subtopic, research, testing, and analysis should be conducted to demonstrate technical feasibility during Phase I and show a path towards Phase II hardware demonstrations. Phase I feasibility studies should also address cost and risk associated with the hot structures technology. At the completion of the Phase I project, in addition to the final report, deliverables should include at least one of the following to aid assessment of technical feasibility: (a) coupons appropriate for thermal and/or mechanical material property tests, (b) arc-jet test specimens, or (c) a subscale nozzle extension test article or analog component. Emphasis should be placed on the delivery of manufacturing demonstration units for NASA testing at the completion of the Phase II contract. In addition, Phase II studies should address scale-up and integration with vehicles that could make use of the developed technology.

Hot structure technology is relevant to the Human Exploration and Operations Mission Directorate (HEOMD), where the technology can be infused into spacecraft and launch vehicle applications. Such technology should provide either improved performance or enable advanced missions requiring re-usability, increased damage tolerance and the durability to withstand long-term space exploration missions. The ability to allow for delivery of larger payloads to various space destinations, such as the lunar south pole, is also of great interest.

The Advanced Exploration Systems (AES) Program would be ideal for further funding a prototype hot structure system and technology demonstration effort. Commercial Space programs, such as Commercial Orbital Transportation Services (COTS), Commercial Lunar Payload Services (CLPS), and Next Space Technologies for Exploration Partnerships (NextSTEP), are also interested in this technology for flight vehicles. Additionally, NASA HEOMD programs that could use this technology include the Space Launch System (SLS) and the Human Landing System (HLS) for propulsion applications.

Potential NASA users of this technology exist for a variety of propulsion systems, including the following:

• Upper stage engine systems, such as those for the Space Launch System,
• In-space propulsion systems, including nuclear thermal propulsion systems,
• Lunar/Mars lander descent/ascent propulsion systems,
• Solid motor systems, including those for primary propulsion, hot gas valve applications, and small separation and/or attitude-control systems, and
• Propulsion systems for the Commercial Space industry, which is supporting NASA efforts.

Finally, the U.S. Air Force is interested in such technology for its Evolved Expendable Launch Vehicle (EELV), ballistic missile and hypersonic vehicle programs. Other non-NASA users include the U.S. Navy, the U.S. Army, the Missile Defense Agency (MDA) and the Defense Advanced Research Projects Agency (DARPA). The subject technology can be both enhancing to systems already in use or under development, as well as enabling for applications that may not be feasible without further advancements in high temperature composite technology.

References

Hypersonic Hot Structures:

Glass, David. "Ceramic matrix composite (CMC) thermal protection systems (TPS) and hot structures for hypersonic vehicles." 15th AIAA International Space Planes and Hypersonic Systems and Technologies Conference. 2008.

Walker, Sandra P., et al. "A Multifunctional Hot Structure Heat Shield Concept for Planetary Entry." 20th AIAA International Space Planes and Hypersonic Systems and Technologies Conference. 2015.

Liquid Rocket Propulsion systems:

“Carbon-Carbon Nozzle Extension Development in Support of In-Space and Upper-Stage Liquid Rocket Engines” paper; Paul R. Gradl and Peter G. Valentine; 53rd AIAA/SAE/ASEE Joint Propulsion Conference, Atlanta, GA; AIAA-2017-5064; July 2017; https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170008949.pdf.

“Carbon-Carbon Nozzle Extension Development in Support of In-Space and Upper Stage Liquid Rocket Engines” presentation charts; Paul R. Gradl and Peter G. Valentine; 53rd AIAA/SAE/ASEE Joint Propulsion Conference, Atlanta, GA; AIAA-2017-5064; July 2017; https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170008945.pdf.

Note: The above references are open literature references. Other references exist regarding this technology, but they are International Traffic in Arms Regulations (ITAR) restricted. Numerous online references exist for the subject technology and projects/applications noted, both foreign and domestic.

Expected TRL or TRL range at completion of the project: 2 to 4

Desired Deliverables of Phase II

Prototypes or components suitable for testing by NASA or Commercial Space partners.

Desired Deliverables Description

At the completion of Phase I project deliverables should include at least one of the following:  coupons appropriate for thermal/mechanical material property tests, arc-jet test specimens, or a subscale nozzle extension test article. Emphasis should be on the delivery of manufacturing demonstration units, with representative structural features, for NASA testing at the completion of the Phase II contract.

State of the Art and Critical Gaps

The current state of the art for composite hot structure components is limited primarily to applications with maximum use temperatures in the 1093° – 1600°C (2000° – 2912°F) range. While short excursions to higher temperatures are possible, considerable degradation may occur. Reusability is limited and may require considerable inspection before reuse. Critical gaps or technology needs include:  (a) increasing operating temperatures to 1700° – 2204+°C (3092° – 4000+°F); (b) increasing resistance to environmental attack (primarily oxidation); (c) increasing manufacturing technology capabilities to improve reliability, repeatability and quality control; (d) increasing durability/toughness and interlaminar mechanical properties (or introducing 3D architectures); and (e) decreasing overall manufacturing time required.

As an alternative to composites, metallic hot structures may reduce operating temperature requirements to near 1000°C (1832°F) in some applications, while offering greater structural reliability, and should also be pursued. Unfortunately advancements in high temperature metals have been a significant gap.

Relevance / Science Traceability

Hot structure technology is relevant to the Human Exploration and Operations Mission Directorate (HEOMD), where the technology can be infused into spacecraft and launch vehicle applications. Such technology should provide either improved performance or enable advanced missions requiring reusability, increased damage tolerance and the durability to withstand long-term space exploration missions. The ability to allow for delivery of larger payloads to various space destinations, such as the lunar south pole, is also of great interest.

The Advanced Exploration Systems (AES) Program would be ideal for further funding a prototype hot structure system and technology demonstration effort. Commercial Space programs, such as Commercial Orbital Transportation Services (COTS), Commercial Lunar Payload Services (CLPS), and Next Space Technologies for Exploration Partnerships (NextSTEP), also are interested in this technology for flight vehicles. Additionally, NASA HEOMD programs that could use this technology include the Space Launch System (SLS) and the Human Landing System (HLS) for propulsion applications.

Potential NASA users of this technology exist for a variety of propulsion systems, including the following:

• Upper stage engine systems, such as those for the Space Launch System,
• In-space propulsion systems, including nuclear thermal propulsion systems,
• Lunar/Mars lander descent/ascent propulsion systems,
• Solid motor systems, including those for primary propulsion, hot gas valve applications and small separation and/or attitude-control systems, and
• Propulsion systems for the Commercial Space industry, which is supporting NASA efforts.

Finally, the U.S. Air Force is interested in such technology for its Evolved Expendable Launch Vehicle (EELV), ballistic missile, and hypersonic vehicle programs. Other non-NASA users include the U.S. Navy, the U.S. Army, the Missile Defense Agency (MDA), and the Defense Advanced Research Projects Agency (DARPA). The subject technology can be both enhancing to systems already in use or under development, as well as enabling for applications that may not be feasible without further advancements in high temperature composite technology.

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Scope Title
Deployable Aerodynamic Decelerator Technology

Scope Description

Background: NASA is advancing deployable aerodynamic decelerators to enhance and enable robotic and human space missions. Applications include Mars, Venus, Titan, as well as payload return to Earth from Low Earth Orbit. The benefit of deployable decelerators is that the entry vehicle structure and thermal protection system is not constrained by the launch vehicle shroud. It has the flexibility to more efficiently use the available shroud volume, and can be packed into a much smaller volume for Earth departure, addressing potential constraints for payloads sharing a launch vehicle. For Mars, this technology enables delivery of very large (20 metric tons or more) usable payload, which may be needed to support human exploration. The technology also allows for reduced cost access to space by enabling the recovery of launch vehicle assets. This subtopic area solicits innovative technology solutions applicable to deployable entry concepts. Specific technology development areas include:

1) Advancements in textile manufacturing technologies that can be used to simplify production, reduce the mass, or reduce the stowed volume of mechanically deployed structures, inflatable structures, or their flexible thermal protection system. Thermal protection concepts can also lead to improvements in thermal management efficiency of radiant and conductive heat transport at elevated temperatures (exceeding 1200 C). Concepts can be either passive or active dissipation approaches. For smaller scale inflatable systems, less than 1.5 meters in diameter, thin-ply or thin-film manufacturing approaches that can be used to reduce the minimum design gauge are of particular interest for inflatable structures. Focus of Phase 1 development can be subscale manufacturing demonstrations that demonstrate proof of concept and lead to Phase 2 manufacturing scale-up for applications related to Mars entry, Earth return, launch asset recovery, or the emergent small satellite community.

2) Concepts designed to augment the drag or provide guidance control for any class of entry vehicle. Concepts can be either deployable or rigid design systems that are suitable to deployable vehicle designs, including methods that modulate vehicle symmetry or adjust lift for active flight control to improve landing accuracy. Designs that decrease the ballistic coefficient by a factor of two to three times are to be considered. Of particular interest are concepts that can be used to modulate the lift or drag of a vehicle for enhanced control. Phase I proof of concept and preliminary design efforts that will lead to, or can be integrated into, flight demonstration prototypes in a Phase 2 effort are of interest.

3) High temperature capable structural elements to support mechanically deployable decelerators that surpass the performance capability of metallic ribs, joints, and struts. Anticipated systems would include composite elements or hybrid approaches that combine metallic structures with high temperature capable interface materials to improve thermal performance. Phase 1 development can be subscale component demonstrations that lead to Phase 2 scale-up and testing in relevant environments.

References

Hughes, S. J., et al, “Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview,” AIAA Paper 2011-2524

Bose, D. M, et al, “The Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Mission Applications Study,” AIAA Paper 2013-1389

Hollis, B. R., “Boundary-Layer Transition and Surface Heating Measurements on a Hypersonic Inflatable Aerodynamic Decelerator with Simulated Flexible TPS,” AIAA Paper 2017-3122

Cassell, A., et al, “ADEPT, A Mechanically Deployable Re-Entry Vehicle System, Enabling Interplanetary CubeSat and Small Satellite Missions,” SSC18-XII-08, 32nd Annual AIAA/USU Conference on Small Satellites

Expected TRL or TRL range at completion of the project: 1 to 4

Desired Deliverables of Phase II

Prototype, Analysis, Hardware, Software, Research

Desired Deliverables Description

Subscale manufacturing demonstration articles for Phase I that can lead to Phase II manufacturing scale up.

State of the Art and Critical Gaps

The current state of the art for deployable aerodynamic decelerators is limited due to novelty of this technology. Developing more efficient, lighter, and thinner flexible thermal protection system component materials with higher temperature capability could potentially enable more efficient designs and extend the maximum range of use of the concepts. Development of efficient guidance control and drag enhancements concepts for deployable vehicles is enabling technology. Novel and innovative high temperature structural concepts are needed for the mechanically deployed decelerator.

Relevance / Science Traceability

NASA needs advanced deployable aerodynamic decelerators to enhance and enable robotic and human space missions. Applications include Mars, Venus, Titan, as well as payload return to Earth from Low Earth Orbit. HEOMD (Human Exploration and Operations Mission Directorate), STMD (Space Technology Mission Directorate), and SMD (Science Mission Directorate) can benefit from this technology for various exploration missions.

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Scope Title

3D Weaving Diagnostics for Validation of Uniform Weaving Processes

Scope Description

NASA is utilizing 3D woven materials to develop Woven Thermal Protection Systems (W-TPS). Examples of recent 3D woven Thermal Protection Systems (TPS) projects include: 3D Multifunctional Ablative TPS (3D-MAT) for compression pads on Orion, Adaptive Deployable Entry Placement Technology (ADEPT) looking at a mechanically deployable aeroshell (similar to an umbrella) that utilizes 3D woven carbon fabric between the ribs, and Heatshield for Extreme Entry Environment Technology (HEEET), containing dual-layer 3D weaves to provide mass efficient TPS solutions for extreme entry environment missions such as to Venus, Saturn and the outer planets. The specialized equipment used to weave 3D woven preforms is based on standard textile equipment that is substantially modified to allow hundreds of layers to be interwoven together. As these complex woven structures are scaled up, it is critical to understand the dynamics of the 3D weaving equipment/hardware and how interactions between different components affect the unit cell of the woven structure and ultimately the material properties.

This subtopic area solicits innovative technology solutions applicable to 3-D woven materials. Specific technology development areas include:

1. Advancements in the understanding of the impact of weaving parameters on the properties of the final weave itself. Looking at developing methods to associate measured weave diagnostics (such as warp tension and beat up force) to understand the effects of woven material parameters (such as fiber volume fraction and yarn crimp), to develop tools to predict the impacts of changes in weaving parameters on final material properties (such as stiffness and strength).
2. Understand what damage may be introduced into the yarns during the weaving operation and the impact of that damage on material performance (such as strength). Objective is to further improve the understanding of how/if key aspects/parameters in the weaving operation (warp tension, beat up force, warp or fill yarns per inch) lead to damage of the yarns and develop methods to reduce weaving damage and/or guidelines to reduce the level of damage induced in the yarns.

References

More info for 3D-MAT, ADEPT, HEEET can be found at: https://gameon.nasa.gov/publications/

Expected TRL or TRL range at completion of the project: 3 to 6

Desired Deliverables of Phase II

Prototype, Analysis

Desired Deliverables Description

Phase I: Assessment study of potential diagnostic techniques
Phase II: Prototype instrument demonstration on a weaving machine demonstrating increased control capability

State of the Art and Critical Gaps

NASA is investing in woven thermal protection systems, both rigid and mechanically deployable, which both come from a 3D weave. The mechanical/structural properties of these weaves are a strong function of nuances in the resultant weave microstructure; nuances such as fiber volume fraction and the level of crimp in warp versus weft direction or damage induced in the yarns during weaving. An enhanced understanding of the effects of the weaving operation parameters on the final weave itself would better enable scale-up of weaving processes (thickness and width) and tailoring of weaves to meet specific mission needs (how does a change in warp tension to reduce fiber volume fraction manifest itself in changes to crimp or other parameters). There is also value in understanding if/where the weaving operation induces damage into the yarn and its impact on material properties. The current state of the art is very empirical for understanding the effects of weaving parameters on material performance/damage. For example, it is recognized that increasing crimp can decrease stiffness in a material, but there are not good tools to predict the impacts of changes in weave parameters (such as warp tension) are on the crimp level in a weave and how that will impact the properties of the final material. This makes it difficult to predict the impacts of changes in weave on properties and understand how sensitive the relationships are. The end result is that this lack of knowledge limits the flexibility end users have, and requires substantial amounts of testing to understand if a given change is important or not. 

Relevance / Science Traceability

Several potential future missions, outlined in decadal surveys, crewed exploration mission studies, and other supporting analyses, have Entry and Descent (ED)/ Entry, Descent and Landing (EDL) architectures: Mars sample return, high speed crewed return, high mass Mars landers, Venus and gas/ice giant probes. With few exceptions, entry vehicle TPS (Thermal Protection System) for these missions will be composed of materials currently under development and without certification heritage.

NASA planetary exploration programs supporting ED/EDL missions are the intended beneficiaries of this subtopic.

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